NASA/CR—2003—212525
`
`
`
`U1tra—High Bypass Engine Aeroacoustic Study
`
`Philip R. Gliebe and Bangalore A. Ianardan
`GE Aircraft Engines, Cincinnati, Ohio
`
`October 2003
`
`GE_1018_001
`
`UTC-2009.001
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`GE V. UTC‘
`
`Trial IPR2016-00952
`
`

`
`The NASA STI Program Office . . . in Profile
`
`Since its founding, NASA has been dedicated to
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`
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`
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`For more information about the NASA STI
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`E-mail your question via the Internet to
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`GE-1018.002
`
`UTC-2009.002
`
`

`
`NASA / CR—2003-212525
`
`
`
`U1tra—High Bypass Engine Aeroacoustic Study
`
`Philip R. Gliebe and Bangalore A. Ianardan
`GE Aircraft Engines, Cincinnati, Ohio
`
`Prepared under Contract NAS3—25269, Task Order 4
`
`National Aeronautics and
`
`Space Administration
`
`Glenn Research Center
`
`October 2003
`
`GE_1018_003
`
`UTC-2009.003
`
`

`
`Acknowledgments
`
`The following people contributed substantially to the study reported herein. Christopher J. Smith provided
`integration of the cycle analysis, flowpath design, mission analysis, and carried out the DOC analyses.
`Paul Feig and Valerie McKay provided the mission analysis. Larry Dunbar and Michael Salay carried out the
`engine flowpath designs. Rick Donaldson, Mark Wagner, and Charlotte Salay provided the engine cycle analyses.
`Dr. Bangalore Janardan and George Kontos provided the engine system noise predictions.
`
`Trade names or manufacturers’ names are used in this report for
`identification only. This usage does not constitute an official
`endorsement, either expressed or implied, by the National
`Aeronautics and Space Administration.
`
`Contents were reproduced from the best available copy
`as provided by the authors.
`
`Note that at the time of research, the NASA Lewis Research Center
`was undergoing a name change to the
`NASA John H. Glenn Research Center at Lewis Field.
`Both names may appear in this report.
`
`NASA Center for Aerospace Information
`7121 Standard Drive
`Hanover, MD 21076
`
`National Technical Information Service
`5285 Port Royal Road
`Springfield, VA 22100
`
`Available from
`
`Available electronically at http://gltrs.grc.nasa.gov
`
`GE-1018.004
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`

`
`Preface
`
`
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`This report was delivered to NASA as an informal document. There were three engine
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`noise studies done by the Allison Engine Company (now Rolls Royce), General Electric
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`Aircraft Engines and Pratt & Whitney in preparation for the Advanced Subsonic
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`Technology (AST) Noise Reduction Program. The objectives of the studies were to
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`identify engine noise reduction technologies to help prioritize the research that was
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`subsequently done by the AST Program. The reports also summarize the predicted
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`performance and economic impact of the noise reduction technologies.
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`The emphasis of commercial turbofan research during the early l990’s was on higher
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`bypass ratio engines. While the technology insertion into service has been slower than
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`expected, many of the results from these studies will remain valid for a long period of
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`time and should not be forgotten by the aerospace community. In 2003, NASA decided
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`to publish all three studies as Contractor Reports to provide references for future work.
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`The quality of the reproduction of the original report may be poor in some sections.
`
`
`
`Dennis L. Huff
`
`
`
`
`
`
`Chief, Acoustics Branch
`NASA Glenn Research Center
`
`
`
`
`
`
`NASA/CR—2003—2l2525
`NASA/CR—2003-212525
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`ULTRA-HIGH BYPASS ENGINE AEROACOUSTIC STUDY
`
`Final Report Prepared for
`
`National Aeronautics and Space Administration
`Lewis Research Center
`Contract NAS3 25269
`Task Order 4
`
`by
`
`Philip R. Gliebe
`and
`
`Bangalore A. Janardan
`
`GE Aircraft Engines
`Advanced Engineering Programs Department
`
`July 8,
`
`l993
`
`ABSTRACT
`
`to identify potential advanced aircraft engine
`A system study was carried out
`concepts and cycles which would be capable of achieving a
`5
`to 10 EPNd8
`reduction in community noise level
`relative to current FAR36 Stage 3 limits
`for a typical
`large-capacity commercial
`transport aircraft.
`The study was
`directed toward large twin—engine aircraft applications in the 400,000 to
`500,000 pound take-off gross weight class.
`
`Four single—rotation fan engine designs were evaluated, over a range of fan
`pressure ratios from 1.3 to 1.75.
`An advanced core design technology was
`assumed, compatible with what can probably be demonstrated by year 2005,
`in
`terms of overall cycle pressure ratio and turbine inlet
`temperature.
`In
`addition,
`two counter-rotating (CR)
`fan engine configurations were studied.
`One of these employed a front-drive, geared fan, and the other was configured
`with an aft-mounted,
`turbine-driven (direct-drive) fan, similar in concept
`to
`the GEAE-developed UDF Engine. Utilizing GEAE design methods, models and
`computer codes,
`the engine performance, weight, manufacturing cost,
`maintenance cost, direct operating cost
`(DOC) and community noise levels were
`estimated for these advanced, ultra-high bypass engine designs.
`
`that significant noise level
`The results obtained from this study suggest
`reductions can potentially be achieved by designing an engine with a
`fan
`pressure ratio of 1.5 or less.
`Selecting fan pressure ratio significantly
`less than 1.5, however, while yielding greater sideline noise reductions,
`provides only small noise reductions at reduced power (cutback and approach),
`while adding significantly to the weight and DOC of the system.
`Significant
`noise reductions were also forecast for the counter-rotating fan engines.
`The
`front-drive, gear-driven CR fan engine, designed for a fan pressure ratio of
`1.3, had significant weight and D.0.C. penalties relative to the
`single-rotation (SR)
`fan counterpart, although the noise levels were 1
`to 2
`EPNd8 lower.
`The rear-drive,
`turbine-driven CR fan, however, was forecast
`
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`lower noise levels relative to its
`as
`to have lower DOC as well
`single-rotation (1.6 fan pressure ratio) counterpart, with very small weight
`penalty.
`
`In summary, several aircraft engine configurations were identified which, with
`further technology development, could achieve the objective of 5 to 10 EPNdB
`reduction relative to FAR36 Stage 3 community noise certification limits.
`Optimum design fan pressure ratio is concluded to be in the range of 1.4 to
`1.55 for best noise reduction with acceptable weight and DOC penalties.
`Further
`in-depth studies in this pressure ratio range are recommended to
`define the best engine architecture in terms of single— vs. counter-rotation,
`geared vs. direct drive fan,
`and separate flow vs. mixed flow exhaust.
`
`NASA"C‘R—2003-212525
`
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`GE-1 01 8.007
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`UTC-2009.007
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`

`
`ULTRA-HIGH BYPASS ENGINE AEROACOUSTIC STUDY
`
`INTRODUCTION
`
`The projected growth of commercial aircraft operations suggests that air
`traffic and passenger-miles will
`increase significantly in the coming decades.
`Many airport operators and rule-making organizations feel
`that
`the current
`FAR36 Stage 3 community noise limits may not be sufficiently stringent
`to
`preclude significant community annoyance around airports.
`Several
`rule-tightening scenarios have been proposed,
`including reducing the current
`FAR36 Stage 3
`limits by anywhere from 3
`to as much as
`10 EPNdB at each
`monitoring point.
`
`Local airports have already imposed their own restrictions to implement noise
`abatement
`in surrounding communities.
`These include night—time curfews,
`night-time operating limits based on certificated noise levels,
`frequency-of-operation restrictions based on noise levels, and landing fees
`based on noise levels.
`
`These local airport noise restrictions are usually more stringent than the FAR
`Stage 3
`limits in terms of equivalent EPNL, although they may be based on
`other metrics such as dBA (Washington National Airport), SENEL (Orange County
`John Wayne Airport),
`and contour area (London Heathrow and Gatwick Airports).
`These local airport restrictions are typically 3
`to 7 dB more stringent when
`cast in terms of equivalent FAR36 EPNL.
`
`increasing rule stringency and the projected
`Given the current climate for
`growth in commercial air traffic,
`it
`is reasonable to expect that noise level
`limits will become significantly lower
`in the next 10 to 20 years.
`The
`current
`technology available to accomplish significant reductions in engine
`noise will
`impose serious performance and/or weight penalties to the
`engine/aircraft system, since all of the known practical methods for reducing
`engine noise have been incorporated in modern high bypass engine designs, at
`least
`to the extent possible within the guidelines of practicality and
`economic viability.
`
`transport
`Engine configurations being considered for future large civil
`aircraft include so-called Ultra-High Bypass
`(UHB) engine cycles, with bypass
`ratios exceeding 10 to 15:1.
`The advantage of a UHB cycle is the significant
`improvement
`in propulsive efficiency and corresponding specific fuel
`consumption that can potentially be attained.
`A significant factor in
`assessing the potential benefit of
`a UHB engine is the achievable core
`technology that can be incorporated, specifically the overall pressure ratio
`(OPR),
`the compressor exit temperature (T3) and the turbine inlet temperature
`(T41).
`
`An important factor in the selection of a new engine cycle and architecture is
`the noise reduction potential, and how much of any identified noise goal needs
`to be achieved by advances in noise reduction and suppression technology vs.
`the "natural" noise reduction which might be achieved from the proper cycle
`selection.
`A proper study is therefore required to assess the potential
`improvements
`in engine performance, weight, cost, complexity, mission
`
`}LkS}tCTL——2003-212525
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`
`focus of the
`The
`including noise.
`economics and environmental emissions,
`present study was to address noise reduction. but
`to provide realistic engine
`concept architectures with reasonable performance and economic assessments, so
`that potential
`low-noise engine concepts could be identified that hold
`promise.
`
`OBJECTIVES
`
`The major objective of the present study was to identify candidate Ultra-High
`Bypass
`(UHB) engine concepts which provide the best noise reduction
`opportunities with the least economic penalties.
`A second objective was
`to
`quantify the effect of bypass ratio (BPR) selection on the acoustics and
`economics of advanced UHB engine concepts.
`A final objective was to identify
`the noise reduction technology improvements required for the best of
`the
`configurations studied,
`and recommend a
`follow-up study and exp’
`‘mental
`development program.
`
`The community noise goal selected for assessing the relative merits of the
`study engines was that the community noise levels should be at
`least 5 to 10
`EPNdB lower than the current FAR36 Stage 3 limits.
`
`SCOPE OF STUDY
`
`large twin-engine civil aircraft application,
`The present study focused on a
`with a 3000 nautical mile mission range and a 250 to 300 passenger capacity,
`similar to the current Boeing 8767-300 and Airbus A300-600 aircraft in service
`today.
`Several advanced engine cycle concepts were selected for evaluation.
`Four single rotation fan engine designs were selected, with design fan
`pressure ratios (FPR) of 1.3, 1.45, 1.6 and 1.75.
`These engines were assumed
`to all have the same core technology.
`i.e.,
`they all had the same overall
`pressure ratio, compressor exit
`temperature and high-pressure turbine inlet
`temperature design points.
`a
`The first was
`Two counter-rotating fan configurations were also studied.
`front-mounted, gear-driven CR fan with a design fan pressure ratio of 1.3.
`This engine concept
`is a Counter-rotating alternative to the 1.3 FPR
`single-rotation fan.
`A second CR fan engine was evaluated which had an
`aft-mounted, direct, CR turbine-driven fan, similar in concept
`to the
`GEAE-developed Unducted fan engine or UDF.
`It was configured with a design
`fan pressure ratio of 1.6,
`and served as
`the CR alternative to the 1.6 FPR
`single-rotation fan engine.
`
`a preliminary design analysis was carried out,
`these six engines,
`For
`consisting of the following steps:
`
`1. Cycle and Engine Architectura Selection
`
`2.
`
`Engine Flowpath Design
`
`3. Engine Cycle Performance Mapping
`
`4. Engine/Aircraft Mission Analysis
`
`}LAS}rCTL——2003-212525
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`5. Community Noise Analysis
`
`6. Noise Reduction Feature Assessment
`
`the basic advanced engines studied,
`for each of
`Step 6 consisted of,
`identifying design changes which would reduce the noise,
`and then evaluating
`the performance, weight,
`and economic impact of these changes and the
`resulting noise reduction benefit.
`For all engines studied,
`the aircraft was
`assumed fixed in size and weight,
`and no advanced aircraft performance
`improvements were assumed.
`
`The baseline selected for referencing all performance, weight, economic and
`noise benefits was an updated version of the Energy Efficient Engine (EEE)
`developed by GEAE under NASA contract
`in the early 1980's,
`reference 1.
`This
`engine, considered to be a current
`technology state-of-the-art demonstrated
`design, was also used as a reference baseline for advanced concept engine
`studies reported in reference 2.
`
`TECHNOLOGY LEVEL ASSUMPTIONS
`
`to select
`The guideline for establishing technology levels for this study was
`what could potentially be available for year 2005 entry into service.
`Based
`on GEAE experience and expertise,
`the following engine technology level
`assumptions were made:
`0
`Compressor Exit Temperature (T3)
`
`— 1390 deg.F
`
`0
`
`0
`
`0
`
`HP Turbine Inlet Temperature (T41)
`
`- 2800 deg.F
`
`Maximum Overall Pressure Ratio (OPR)
`-
`High Pressure Compressor (HPC)
`-
`Fan + Booster (LPC)
`
`- 55:1
`— 27:1
`- 2.04:1
`
`Component efficiencies - based on a 5 percent reduction in losses
`relative to current technology.
`
`As mentioned in the previous section, current state-of-the-art aircraft
`performance was assumed.
`
`ENGINE CYCLE SELECTION PHILOSOPHY
`
`it was the intent of this study to evaluate
`As discussed in the introduction,
`the effect of increasing bypass ratio on community noise.
`From the standpoint
`of engine cycle selection,
`for a given thrust requirement,
`the bypass ratio is
`a product of the fan pressure ratio selected and the core technology level
`(OPR and T41) assumed. Also,
`from a noise reduction point-of-view,
`the
`exhaust jet mixing noise, a primary contributor at full power for current high
`bypass engines,
`is dictated to a great extent by the fan pressure ratio.
`The
`FPR selected sets the fan jet exhaust velocity, which in turn sets the jet
`exhaust noise level, since jet noise is roughly proportional
`to the sixth
`power of the jet velocity for a given thrust.
`
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`Fan pressure ratio was therefore selected as the major independent variable to
`be studied. and the bypass ratio was considered as a computed result based on
`the core technology assumed and the thrust requirement.
`FPR was varied from
`1.75,
`typical of current high bypass engines (but now with an advanced core),
`down to 1.3.
`FPR values less than 1.3 were felt to be impractical
`in that the
`resulting fan and nacelle size would not be compatible with an under-the-wing
`installation on a
`typical B767/A300 type aircraft.
`The study resources
`limited the number of FPR values to four cycles: 1.3, 1.45 ,1.6, and 1.75.
`
`it was
`for example),
`From previous preliminary design studies (reference 2,
`concluded that
`the imcompatibility between fan and Low Pressure Turbine (LPT)
`speed for achieving good component efficiencies and low number of LPT stages
`as FPR is reduced implies that
`a geared fan should be used for FPR values
`significantly less than about 1.5.
`The FPR=l.45 and 1.3 engine cycle
`architectures were therefore designed as gear-driven fan engines.
`
`study engines
`A mixed flow exhaust system architecture was assumed for all
`with FPR of 1.45 and higher.
`It was felt that this would result
`in better
`Specific Fuel Consumption (SFC),
`and lower noise.
`Separate Flow exhaust
`system architecture was assumed for the FPR - 1.3 engines, because it was felt
`that the large nacelle size required at this low fan pressure ratio would make
`a mixed flow system much too heavy and yield high nacelle drag because of the
`much larger wetted area.
`
`It was of interest to evaluate whether a counter-rotating fan offered a noise
`reduction advantage relative to a single-rotation fan.
`conceptually, having
`two rotors produce the same total FPR as one rotor would allow the two rotors
`to run at lower tip speeds, and therefore potentially produce less total noise
`than one rotor producing the same FPR at a significantly higher tip speed.
`
`For reasons to be
`a gear-driven CR fan seemed the best approach.
`For low FPR,
`discussed later,
`the use of a gearbox for counter-rotation imposed severe
`restrictions on the speed ratio,
`torque ratio and fan exit swirl,
`and the
`maximum reasonable FPR that gave a sensible engine was found to be 1.3.
`A 1.3
`FPR engine with a CR gear-driven fan was therefore selected for evaluation.
`
`To avoid having to fit two fan
`A direct-drive CR fan was also evaluated.
`shafts through the middle of a two-spool core, with all
`the conflicting
`requirements for bearings, shaft sizes,
`and core flow path constraints,
`a
`rear-mounted fan was selected.
`A higher FPR of 1.6 was selected,
`to take
`advantage of the reduced tip speed requirement at higher FPR, and potentially
`provide a quieter engine at a smaller fan diameter. This engine is similar in
`concept
`to the GEAE-developed UDF engine, as discussed in reference 3, but
`with a ducted CR fan and much lower bypass ratio.
`
`Table 1 lists the primary cycle and geometry parameters for the engine
`configurations selected for study.
`The selection in some cases involved some
`iterations to arrive at an engine cycle and engine architecture that was
`reasonable,
`in the sense that there were no known barrier problems that needed
`to be overcome to make the engine viable.
`Table 2 summarizes the component
`efficiencies that were assumed for each configuration.
`
`the decisions for which engines should be
`It should be understood that
`gear-driven vs. direct drive and which engines should be mixed flow vs.
`separate flow were based on prior experience with preliminary design study
`
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`results, and probably need verification if a final engine design concept were
`to be pursued.
`A thorough design optimization study would be required to more
`carefully weigh the trades between performance, noise, weight, cost,
`complexity, maintainability and customer acceptance, before deciding on the
`fan drive and exhaust system architecture.
`
`SINGLE-ROTATION FAN ENGINE DESIGNS:
`
`FLOW PATH DESIGN SUMMARY
`
`The advanced engine preliminary designs were generated using a GEAE computer
`code called FLOHPATH.
`This code utilizes GEAE modelling experience for
`component aerodynamic performance, mechanical design,
`and manufacturing
`(material selection, costs, etc.).
`For a given cycle,
`the code FLOHPATH will
`define an entire engine, using appropriate mission requirement data for each
`component.
`The engine overall and component dimensions are estimated. and all
`part weights are determined,
`including blade and vane airfoils, disks,
`frame
`structures, bearings, seals, shafts, and controls and accessories.
`A typical
`subsonic mission engine FLOHPATH output
`is shown in figure 1.
`
`The advanced engine designs selected for study were generated using the
`FLOHPATH code.
`Figure 2 shows the FLOHPATH generated engine cross-section for
`the baseline updated EEE engine.
`This engine, described in references 1 and
`2,
`serves as the reference for the performance, noise, weight and D.0.C.
`assessments for the advanced engines.
`In its original
`form (reference 1),
`it
`was built and tested,
`and GEAE has evaluated its performance and noise
`characteristics.
`
`The engine FLOHPATH cross-sections for the two direct-drive single-rotation
`engines, Engine 1
`(FPR-1.75) and Engine 2
`(FPR-1.6), are shown in figure 3.
`The FLOHPATH cross-sections for the two gear-driven single-rotation engines,
`Engine 3
`(FPR-1.45) and Engine 4 (FPR=1.3), are shown in figure 4. Note that
`Engines 1, 2, and 3 all have mixed-flow exhaust systems.
`It
`is also
`noteworthy that
`the HP compressor has fewer stages (8 vs. 10)
`for the geared
`engines,
`and one HP turbine stage for the geared engines vs.
`two for the
`direct-drive engines.
`The gear-driven fan engines are therefore shorter from
`fan rotor exit to turbine rear frame exit.
`These single-rotation fan engines
`are sometimes labelled as S75, S60, S45, and $30 for Engines 1, 2, 3, and 4,
`respectively.
`
`It can be seen from figures 3 and 4 that all of the advanced study engines
`employ an integral vane/frame outlet guide vane (OGV) design for the fan.
`This provides a larger axial spacing between the fan rotor and the OGV, which
`helps keep the fan interaction-generated tone noise lower than would be the
`case with a separate OGV row in front of the fan frame struts.
`
`FRoN'r-MouNTEn COUNTER-ROTATION FAN DESIGNS:
`
`two
`As discussed in the previous section on Engine Cycle Selection,
`counter-rotation fan engine designs were studied.
`The front mounted.
`gear-driven fan engine was designed for a fan pressure ratio of 1.3. This was
`found to be the about
`the highest
`fan pressure ratio that would still result
`in a reasonable engine configuration.
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`the bore size of the core engine becomes
`For a very high bypass ratio engine,
`quite small, and there is insufficient
`room for two counter-rotating shafts
`for driving the fans directly by an LP counter-rotating turbine.
`Thus,
`a
`single shaft LP turbine was selected with a gearbox to drive the two
`counter-rotating fan rotors.
`A planetary gearbox design was studied, and the
`gearbox constraints dictated the selection of the fan pressure ratio.
`
`to be addressed is that of keeping the rotor inlet
`The first constraint
`relative Mach number at or below unity. This is desirable from the standpoint
`of mimimizing noise.
`As
`long as the fan pressure ratio is low enough,
`this
`constraint
`is easily satisfied on the front
`fan rotor. Because of
`the swirl
`added by the first rotor and the counter-rotating wheel
`speed of the second
`rotor,
`the second rotor will have a higher relative Mach number
`than the
`first, especially at
`the hub.
`The selection than involves an iterative
`process of choosing an overall
`fan pressure ratio. selecting the forward/aft
`rotor pressure ratio split and evaluating the implied rotor tip speeds,
`torque
`ratios, and inlet relative Mach numbers.
`Figure 5
`shows a typical design
`curve used for selecting fan tip speeds as a function of fan pressure ratio.
`
`involved keeping the second rotor exit swirl as small as
`A second constraint
`possible,
`in order to reap the "inherent advantage" of Counter-rotation that
`no OGV row is needed.
`This constraint
`implies keeping the rotor torque ratio
`as close to unity as possible.
`This also helps keep the number of planet
`gears required to a minimum.
`Figure 6 shows the design trends for dependency
`of exit swirl and number of planet gears on fan (front-to-rear) torque ratio.
`However,
`for a CR output shaft, having torque ratio close to unity requires a
`much higher gear ratio,
`so that,
`for a given fan speed,
`the LP turbine must
`run at
`a much higher speed.
`Figure 7
`shows
`the required gear ratio as a
`function of torque ratio.
`
`An additional constraint to consider is that of LP turbine exit flow area and
`
`flow area required as a
`the turbine exit
`Figure 8 shows
`speed combined.
`function of fan pressure ratio. Higher Bypass ratios require greater LP
`turbine expansion and greater exit area to pass the flow.
`The parameter
`combination
`AN2
`or (Exit Area)*(RPM-squared)
`is a measure of the LP turbine
`last stage blade root stress. Design limits on this parameter therefore add a
`constraint
`to the selection of fan pressure ratio and torque ratio, as shown
`in figure 9.
`A limit on "AN2" of 45 was selected as being as high as possible
`without significantly exceeding best available technology and experience.
`
`fan
`a
`through 9 were employed to arrive at
`The trends shown in figures 5
`and a
`a speed ratio of 0.8 (aft/forward),
`overall pressure ratio of 1.3,
`the exit swirl down to 7 or 8 degrees.
`the
`torque ratio of 1.5.
`This kept
`number of planet gears down to eight, and the gear ratio down to about 5:1.
`A separate flow exhaust system was also selected for this engine, because the
`bypass ratio was high enough (BPR=l5.75)
`that a mixed flow system would offer
`no significant performance advantage. Also,
`the mixed flow benefit on jet
`noise would be very small, and, as will be discussed later,
`the jet noise
`contribution itself is small at any rate for this engine cycle.
`
`the resulting engine cross-sections as generated by the
`Figure 10 shows
`FLOHPATH program for the front-mounted, gear—driven, counter-rotating fan
`engine.
`Four versions are shown in figure 10, corresponding to four different
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`These
`combinations of fan blade numbers and rotor-to-rotor axial spacings.
`four variants were selected to evaluate the influence of blade number
`selection and axial spacing on community noise, and to determine the economic
`sensitivities to these changes.
`This engine configuration is referred to as
`engine 5, and the four variants are labelled SA, 58, SC, and SD, as summarized
`in table 3. This engine is also sometimes referred to as CF30.
`
`Engine 58 is the baseline from which the other variants (SA, SC, and 50) were
`selected.
`Figure 10a shows a comparison of Engine SA (top) with the baseline
`SB (bottom).
`The difference is the increase in forward-to-aft rotor axial
`spacing-to-chord ratio from 1.2 (SB)
`to 2.5 (SA).
`The advantage of 5A over 58
`was expected to be a reduction in interaction noise, but at
`the expense of
`engine length and weight.
`
`The
`a comparision of Engine SC (top) with SB (bottom).
`Figure 10b shows
`difference is the change in rotor blade numbers from 19 forward rotor blades
`and 15 aft rotor blades (58)
`to 15 forward rotor blades and 19 aft rotor
`blades (SC).
`The intent of
`this variant was
`to produce negative—spinning
`interaction modes, which would have greater transmission loss through the
`forward rotor,
`thus reducing the forward-radiated interaction tone levels.
`In
`addition,
`the number of
`frame vane/struts was
`increased from 36 to 46,
`in
`order to preserve cut-off of the aft rotor BPF (blade-passing frequency)
`tone
`produced by aft rotor wake-strut
`interactions.
`
`This
`a comparision of Engine SD (top) with SB (bottom).
`Figure 10c shows
`Engine is a variant of Engine SC with the rotor-to-rotor axial
`spacing
`increased from 1.2 to 2.5 projected chords.
`This engine is the longest and
`heaviest of the four.
`
`Arr-Mourrrao COUNTER-ROTATING FAN DESIGN:
`
`The final Engine configuration studied is an aft-mounted, counter—rotating fan
`engine design.
`The FLOHPATH-generated engine cross-section is shown in figure
`11.
`This configuration is similar in concept
`to the Unducted Fan Engine,
`reference 3. which has
`a
`two-spool gas generator core which drives a
`free-wheeling, counter-rotating turbine, which in turn powers
`the two,
`counter—rotating fan stages.
`In selecting the cycle for this engine concept,
`gearbox constraints were not
`a consideration, and the fan shafts do not have
`to pass through the core.
`It was
`therefore decided to take advantage of the
`two fan stages and select a fan pressure ratio which was reasonably high,
`so
`as to provide a compact engine, but not
`so high as to produce high jet noise.
`A fan total pressure ratio of 1.6 was selected as being comparable to Engine
`S60,
`its single-rotation counterpart, and potentially would have a propulsive
`efficiency and noise advantage as well.
`
`requires
`The fan nacelle designed for this engine, as shown in figure 11,
`support struts both forward of and behind the fan.
`The forward struts
`potentially could shed wakes into the fan rotors, producing additional noise,
`so a
`large axial
`spacing of 3 strut chords was selected to minimize this
`effect.
`Further,
`the forward strut and first rotor blade numbers were
`selected to provide a high spinning mode number
`(28 - 4 - 24)
`so that
`the
`nacelle treatment between the struts and rotor would have better attenuation
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`
`spacing criterion of 2.5
`the rotor-to-rotor axial
`In addition,
`performance.
`forward rotor projected chords was easily accommodated because the
`forward-to-aft fan power frame spacing was also needed to fit in the required
`number of turbine rotor stages.
`
`Enema WEIGHT AND COST Esunnesz
`
`the engine
`for all of
`Engine weight and cost estimates were made
`configurations described in the above paragraphs, using the FLOHPATH code.
`Figure 12 summarizes the engine-plus—nacelle weights, manufacturing costs. and
`maintenance costs for the four single-rotation and two counter-rotation fan
`engines,
`in terms of percent changes from the baseline EEE values.
`The
`component contributions of the fan, booster, HP compressor, HP turbine, and LP
`turbine systems to the total changes in engine weight, manufacturing cost, and
`maintenance cost are shown in figures 13, 14, and 15, respectively.
`
`The front-driven, counter-rotating fan engine had four variants, as shown in
`figures 10a-c.
`The corresponding variations in weight, manufacturing cost,
`and maintenance cost are shown in figure 16.
`In general,
`the significant
`discriminator is the axial
`spacing difference, as
`shown by the fact
`that
`engines SA and SD have similar weights and costs, and engines SB and 5C have
`similar weights and costs, with engines 53 and SC being slightly lighter and
`cheaper.
`
`For each of the engine cross-sections shown in figures 3 and 4, axial spacing
`and inlet length were evaluated from an acoustic design viewpoint.
`The axial
`spacings and inlet
`lengths were then modified,
`if necessary,
`to provide
`acoustically prudent axial
`spacing/chord rat