Advanced Single-Aisle Transport Propulsion Design Options
`Revisited
`
`Mark D. Guynn*
`NASA Langley Research Center, Hampton, VA, 23681
`
`Jeffrey J. Berton†, Michael T. Tong‡, and William J. Haller§
`NASA Glenn Research Center, Cleveland, OH, 44135
`
`Future propulsion options for advanced single-aisle transports have been investigated in
`a number of previous studies by the authors. These studies have examined the system level
`characteristics of aircraft incorporating ultra-high bypass ratio (UHB) turbofans (direct
`drive and geared) and open rotor engines. During the course of these prior studies, a number
`of potential refinements and enhancements to the analysis methodology and assumptions
`were identified. This paper revisits a previously conducted UHB turbofan fan pressure ratio
`trade study using updated analysis methodology and assumptions. The changes in
`propulsion, airframe, and noise modeling are described and discussed. The impacts of these
`changes are then examined by comparison to the previously reported results. The changes
`incorporated have decreased the optimum fan pressure ratio for minimum fuel consumption
`and reduced the engine design trade-offs between minimizing noise and minimizing fuel
`consumption. Nacelle drag and engine weight are found to be key drivers in determining the
`optimum fan pressure ratio from a fuel efficiency perspective. The revised noise analysis
`results in the study aircraft being 2 to 4 EPNdB (cumulative) quieter due to a variety of
`reasons explained in the paper. With equal core technology assumed, the geared engine
`architecture is found to be as good as or better than the direct drive architecture for most
`parameters investigated. However, the engine ultimately selected for a future advanced
`single-aisle aircraft will depend on factors beyond those considered here.
`
`I. Introduction
`
`S
`
`INCE 2006, NASA has been conducting on-going trade studies to assess propulsion options for an advanced
`single-aisle (Boeing 737/Airbus A320 class) aircraft, initially as part of the Subsonic Fixed Wing (SFW) Project
`and continuing today in the Fixed Wing Project. The focus of these efforts has been to assess potential technology
`paths for reaching the NASA “N+1” subsonic transport system level goals shown in Fig. 1. This multi-year, multi-
`phase activity began with an initial concentration on ultra-high bypass ratio (UHB) geared and direct drive turbofan
`engines. Initial findings, along with multiple interactions with industry partners, were used to refine the UHB
`analysis process and assumptions over a period of a few years. Then the focus shifted to open rotor (OR) engine
`options for this class of vehicle. Following a similar path, the initial results were updated and refined based on
`interaction with industry partners. Numerous technical reports and papers document the results of these studies.1-10
`Continuous improvement through publication and subsequent discussions and interactions has been a defining
`characteristic of this multi-year activity. NASA’s modeling and analysis tools have also been in a state of continual
`development over the course of this activity. The analysis processes and “best practices” have evolved over time.
`Although leading to better study results, such improvements make comparison of recent results to previous findings
`problematic. Because of the evolving methodologies and modeling assumptions, the Fixed Wing Project recently
`initiated an effort to update the previous UHB studies, enabling consistent comparisons with the more recent open
`rotor studies described in Ref. 10.
`
`
`*Aerospace Engineer, Aeronautics Systems Analysis Branch, Mail Stop 442, Senior Member AIAA.
`†Aerospace Engineer, Multidisciplinary Design, Analysis & Optimization Branch, MS 5-11, Senior Member AIAA.
`‡Aerospace Engineer, Multidisciplinary Design, Analysis & Optimization Branch, MS 5-11.
`§Aerospace Engineer, Multidisciplinary Design, Analysis & Optimization Branch, MS 5-11.
`
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` 2013 Aviation Technology, Integration, and Operations Conference
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`Figure 1. NASA subsonic transport system level metrics.
`
`
`Much has changed in the aircraft industry since 2006 as well. At that time, the expectation was that completely
`new 737 and A320 replacement aircraft would be available in the mid-2010s.11 There are two completely new
`vehicles in this size class on the horizon, the Comac C919 and the Irkut MC-21, from China and Russia respectively.
`However, the major western manufacturers, Boeing and Airbus, have chosen to instead re-engine their existing 737
`and A320 models, with entry-into-service of the A320neo projected in 2015 and the 737 MAX in 2017. Although
`the expected all-new Airbus and Boeing aircraft have not emerged, new engines are being developed. When the
`NASA UHB initial feasibility study was started in 2006, the Pratt & Whitney PW1000G and CFM International
`LEAP products were not yet in development. The potential competition between an advanced geared turbofan
`engine and an advanced direct drive turbofan engine was theoretical. With the geared PW1100G and direct drive
`CFM LEAP-1A engines both offered on the A320neo, that competition now exists in the single-aisle engine market.
`Even though the entry-into-service dates for these new engines match the technology dates of the original NASA
`study, the NASA technology assumptions were aggressive and still represent some advancement beyond the current
`LEAP and PW1000G engines. This level of technology, referred to as N+1 in NASA terminology, is 5-10 years
`beyond the new P&W and CFM engines. This is an important distinction since it has not been, nor is it now, the
`authors’ intent to compare or assess engines that are available in today’s market. Rather, the focus is on propulsion
`choices that will need to be made in the future.
`
`II. Study Objectives and Approach
`The objective of this study is to revisit previous analyses to assess advanced geared and direct drive turbofans
`using updated and refined analysis processes and assumptions. The impacts of these changes on prior study
`conclusions are evaluated. Ultimately, the objective is to enable comparison of the fuel burn, noise, and emissions of
`geared turbofan, direct drive turbofan, and open rotor propulsion system options using equivalent technology
`assumptions and a consistent analysis process to permit informed trade-offs among these three options. The general
`approach taken for this study is to develop analytical models of advanced, two-spool ducted turbofan engines,
`combine them with an advanced technology airframe model, design the overall system to meet mission requirements
`and constraints, and assess the resulting noise, fuel consumption, and emission characteristics.
`
`III. Modeling and Analysis Methodology
`
`A. Propulsion Modeling
`Propulsion system modeling is performed using NPSS (Numerical Propulsion System Simulation)12-14 for cycle
`analysis and performance and WATE (Weight Analysis of Turbine Engines)15-17 for aeromechanical design and
`weight/dimension estimates. Estimates for NOX emission indices are obtained from an empirical correlation
`representing an advanced, low NOX combustor. Reference 2 provides more details on this empirical NOX
`correlation, which was developed by NASA combustor technologists during the latter stages of NASA’s Ultra-
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`Efficient Engine Technology Program. All engines in the study are two-spool, separate flow turbofans designed with
`the same Aerodynamic Design Point (ADP) of Mach=0.8 at 35,000 ft and same Overall Pressure Ratio (OPR) of 42
`at the ADP. In addition to investigating geared and direct drive architectures, two different compressor work splits
`are considered. The “low work” engines have a lower pressure rise across the low pressure compressor (LPC) (and a
`higher pressure rise across the high pressure compressor (HPC)) compared to the “high work” engines. Low fan
`pressure ratio (FPR) engine cycles generally require some type of variable geometry for proper operation across the
`flight envelope; the use of a variable area bypass nozzle is assumed in this study. Since the variable area nozzle has
`a weight penalty, it is only applied when necessary to achieve the desired fan surge margins throughout the
`operating envelope. The variation in fan drive approach, compressor work split, and fan pressure ratio results in a
`total of twelve different engines in the study.
`Changes to the engine modeling and assumptions for this updated analysis are summarized in Table 1. The first
`change is new nominal design thrust values resulting from changes in the aircraft modeling discussed in the next
`section. The second change is a correction to the way the bypass ratio is set for each engine design. In engine cycle
`design, the fan pressure ratio is usually treated as an independent variable selected by the designer. With the FPR
`set, the bypass ratio is usually varied to match a preselected target value for the ratio of the core flow and bypass
`flow jet velocities. Generally speaking, when bypass ratio is set in this manner, the same amount of energy is left in
`the core stream as the engine trade space is explored. (Alternatively, an extraction ratio – the ratio of the bypass
`nozzle to core nozzle total pressures – may be held constant to set the bypass ratio.) In Ref. 4, the jet velocities used
`to set bypass ratio were the actual velocities exiting the convergent nozzles. But when the nozzles are supercritical at
`the top-of-climb conditions, they are choked and the nozzle exit velocity is always sonic. Unintentionally, variable
`amounts of energy were therefore left in the core, which led to an inconsistent design space exploration. In the
`current study, this problem is corrected by using nozzle exit velocities ideally expanded to ambient static pressure to
`ensure consistency.
`The definition of the nacelle geometry is also refined for the current assessment. The inlet length-to-diameter
`(L/D) ratio is set by the diffuser exit-to-throat area ratio as described in Ref. 18. The maximum nacelle diameter is
`set by the design values of engine critical mass flow ratio and drag-rise Mach number, using the methodology
`described in Ref. 19. This methodology sets nacelle maximum diameter based on a compromise between reasonable
`nacelle operating margin and low nacelle drag. As shown in Table 1, the new approach results in a nacelle-to-fan
`diameter ratio which is within 1% of that used in Ref. 4.
`
`Table 1. Summary of changes to engine modeling.
`
`Modeling in Reference 4
`
`Current Study
`
`Engine Thrust Sizing
`
`ADP: 5,000 lb
`SL, M=0.25: 17,500 lb
`
`ADP: 5,100 lb
`SL, M=0.25: 18,750 lb
`
`Bypass Ratio
`
`Set by actual jet velocity ratio at ADP
`
`Set by ideal jet velocity ratio at ADP
`
`Inlet L/D ratio
`
`Constant 0.5
`
`Set by diffuser area ratio; ~ 0.4
`
`Nacelle Maximum Diameter 1.23 times fan diameter
`
`1.22 times fan diameter
`
`Variable Area Nozzle
`
`Exit areas not set properly at low
`altitude
`
`Low altitude error corrected
`
`LPT Adiabatic Efficiency
`
`Constant loading; efficiency function
`of LPT cooling level
`
`Efficiency set by number of stages
`and Stewart work-speed parameter
`
`Hot-section Cooling
`
`Gearbox weight
`
`Aggressive cooling effectiveness
`assumptions
`
`Cooling assumptions based on
`current technology large engine
`
`Original WATE correlation
`(developed by Boeing)
`
`NASA updated correlation
`
`
`With a variable area bypass nozzle, the bypass nozzle exit area is normally varied as flight conditions change to
`keep the fan operating at peak efficiency. In performing the current study, it was discovered that in the previous
`study (Ref. 4) a user error in the NPSS solver prevented the bypass nozzle exit areas from automatically increasing
`at lower altitudes. Although correct at mid to high altitudes, the exit areas were somewhat smaller than they should
`have been near sea level. Since the engine cycle was designed using a multiple design point solution, this error
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`impacted the entire cycle design and the performance everywhere, not just a low altitude. Engine noise was also
`affected by the smaller than desired bypass nozzle exit area. The impacts of this error (and the jet velocity ratio error
`noted above) are not great enough to render the previous results invalid, but they are nevertheless corrected in the
`current study.
`In the prior work, the LPT loading was kept constant and the stage count was allowed to grow as needed to
`accommodate the lower spool speeds at low fan pressure ratios. This resulted in an LPT adiabatic efficiency that was
`essentially constant (varying slightly due to different cooling levels), but an excessive number of LPT stages for
`low-FPR, direct drive engines. This approach was used realizing that engines with many LPT stages would be heavy
`and not competitive once evaluated in an airplane sizing calculation. The current assessment uses an iterative
`procedure between the cycle and aeromechanical analyses (i.e., between NPSS and WATE) to determine a more
`appropriate LPT stage count and weight for the direct drive engines. The LPT stage count is now limited to a
`maximum of nine stages. This is achieved by allowing the loadings to vary, within reason, and accepting the
`resulting efficiency penalty. A turbine model relating loading, stage count, and efficiency developed by Warner L.
`Stewart20 is used for the iteration procedure. This model assumes equal stage work, equal mean blade tip speeds, and
`equal stator exit angles. The cycle is initially analyzed assuming three LPT stages and an efficiency of 94%. An
`aeromechanical analysis is then performed to determine the work-speed parameter and estimate a new efficiency
`using the Stewart efficiency model. If the efficiency is less than 92%, additional LPT stages are added until the
`efficiency is greater than 92% or the number of stages reaches the maximum of nine. The cycle analysis is then
`repeated with the new efficiency estimate and the results fed to the aeromechanical analysis to compute a new work-
`speed parameter. This process is iterated until the cycle and aeromechanical analyses are consistent.
`The engine hot-section cooling assumptions were also revised. In Ref. 4, a cooling effectiveness was used to
`determine the cooling bleed flow rates directed to each turbine blade row, given the hot gas temperature and the
`maximum allowable airfoil surface temperature. Upon review, the cooling effectiveness values assumed were too
`aggressive for the N+1 timeframe. In the present study, values of cooling effectiveness appropriate for a current-
`technology GE90 engine are used. The rationale is that engine temperatures and cooling technologies appropriate for
`a current large engine could be transferred to engines in the single-aisle thrust class by the N+1 timeframe. The total
`cooling increased from about 17% to about 19% as a result of this change.
`Finally, an updated weight correlation is used to estimate the weight of the gearbox system for the geared fan
`engines. This empirical correlation, described in Ref. 21, was developed based on weight data from over fifty
`rotorcraft, tiltrotor, and turboprop aircraft. Use of this correlation makes the gearbox weight methodology in this
`study consistent with the single-aisle open rotor study.7-9 The resulting gearbox weights are within 45 lbs of those
`calculated with the older methodology used in Ref 4.
`
`B. Airframe Modeling
`The basic airframe modeling approach is described in detail in Refs. 2 and 4. In general, the same approach is
`retained for the current study. Two areas where enhancements have been made are the baseline modeling and the
`aircraft sizing procedure.
`As described in Ref. 2, the 737-800 (with winglets) is used to develop a baseline analytical model from which
`the benefits of future advanced technologies are assessed. The 737-800-like analytical model used as a starting point
`in the prior studies has been shown to compare favorably with published performance characteristics of the 737-800,
`particularly with regard to fuel consumption.22 However, over the years since the model was initially developed,
`new analysis tools and methods have been incorporated into the standard analysis suite. One example is an
`improvement in the methodology used to size vertical and horizontal tails. Another example is a more explicit
`accounting for propulsion system installation weight. Additionally, the core analysis tool, the Flight Optimization
`System (FLOPS),23 has been modified over the years, which has enabled some enhancements to the modeling “best
`practices.” In order to incorporate these changes in the propulsion trade study, it was necessary to first apply them to
`the baseline vehicle and re-calibrate the model. The NASA model of the CFM56-7B-like engine used on the
`baseline model has also been updated. The baseline engine model used in the prior studies was incomplete and it
`was not possible to generate the data needed to do noise analysis. This required developing a separate CFM56-7B-
`like engine for noise validation. The incomplete engine model has been replaced with a new, complete NPSS model
`to enable a single, consistent model that can be used in multiple types of analyses. There is minimal impact of these
`modeling changes on the 737-800-like starting point, since by design the model is re-calibrated to match the
`published data. However, the new calibration factors are propagated through the rest of the study vehicles and
`impact the results for those vehicles.
`In advanced vehicle concept and technology trade studies, a passenger load and range capability is typically
`specified for the design mission and all the study vehicles are sized to be able to fly that mission (under typical
`
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`reserves and air worthiness constraints). The assumption behind this approach is that it enables a consistent
`comparison of multiple vehicles with the same mission capabilities. In reality, however, the ofi'-design capabilities
`can be quite diflerent. As an aircraft becomes more efficient through advanced technologies, the value of a pound of
`fuel as measured in range capability increases. This means that the decrement in range associated with trading that
`pound of fuel for more payload also increases. As a result, the range capability of the advanced, more efficient
`vehicles can be less for off-design, high payload missions than the baseline vehicle. This loss in capability at high
`payload is undesirable. It is impossible to exactly match the payload-range capability of an advanced vehicle to that
`of the baseline vehicle because of the fimdamental difference in the impact of fuel weight changes on range. It is
`possible, however, to constrain the range capability of the advanced vehicle to be equal to or greater than the
`baseline vehicle and avoid a loss in capability. A new aircraft sizing procedure is used in this study that sets a
`minimum range capability of 2125 nm at maximum payload and matches range capability of 3250 nm at the
`nominal design payload. Maximum takeoff weight, maximum landing weight, and maximum fuel weight are iterated
`until the desired payload-range capabilities are met. The result of this new approach is illustrated by the payload-
`range diagrams shown in Fig. 2, representing the baseline current technology aircraft and the twelve advanced
`vehicles modeled in the study. By imposing range perfonnance constraints at both the maximum payload and design
`payload conditions, the payload-range capabilities are similar for all of the study vehicles.
`Max_Paytoad
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`.:s§:§'§§§§§
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`5oo1ooo150o2ooo25oo3ooo35oo4ooo45oo
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`Range,nm
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`Figure 2. Comparison of payload-range capabilities for study aircrafi.
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`C. Noise Analysis
`Several changes were made to the noise prediction procedures described in Ref. 4. One basic difference is the
`use of ANOPPM” Level 30, which includes several upgrades and bug fixes since Level 26 (the version used in Ref.
`4).
`
`The second difference is the manner in which the takeoff trajectory is modeled. In Ref. 4, the noise abatement
`throttle cutback was assumed to occur at 16,000 fl lrom brake release. Performing the cutback at this location (well
`short of the Part 36 flyover noise monitor location at 21,325 fix) ensured that the flyover eflective perceived noise
`level (EPNL) noise signature consisted of the engines in their throttled, cutback state. However, in noise certification
`practice, the time of the noise abatement throttle cutback is often varied to minimize the flyover EPNL. The throttle
`cutback may be delayed until the airplane is quite near the flyover noise monitor, giving the airplane the opportlmity
`to climb higher and reduce its noise level at the flyover point by spherical spreading effects. The tone-weighted
`perceived noise level (PNLT) vs. time noise history for a minimized flyover EPNL will exhibit two peaks:
`the first
`reflecting the airplane at maximum throttle, gaining altitude quickly but still short of the monitor; and the second at
`reduced throttle, climbing only at a four percent climb gradient, and passing over the monitor. In addition to
`changing the cutback location, a brief acceleration segment at 1000 ft altitude used in Ref. 4 was omitted in favor of
`a constant speed climb. This also allows the airplane to achieve a higher altitude over the flyover monitor and is
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`more representative of a Part 36 certification takeoff. These two modifications to the trajectory calculations result in
`an increase in the altitude over the flyover monitor of over 600 ft. The flyover EPNLs in this study are reduced by
`more than 3 EPNdB relative to the results computed using the procedures in Ref. 4.
`A third difference is the use of a new hardwall fan noise analysis that is more appropriate for advanced fan
`designs likely to be in service in the N+1 timeframe. In Ref. 4, the fan noise was predicted using a method
`developed by General Electric.27 GE’s method consists of a recalibration of the original fan noise method developed
`for ANOPP by Heidmann.28 While the overall structure of Heidmann’s original empirical method remained intact,
`GE adjusted the method’s numerical constants to predict fan noise at levels that reflected GE’s experience base with
`large turbofans in service just prior to 1996: the CF6-80C2, CFM56, E3, and QCSEE engines (see Ref. 27 for
`details). These engines have fans with relatively narrow chords, straight blades, and high pressure ratios; whereas
`modern fans are designed with wider chords, swept and contoured blades, and often have lower pressure ratios and
`tip speeds. In this study, another recalibration of the Heidmann fan noise method is used. In 2006, acoustic
`investigators employed by Diversitech, Inc., working under contract with NASA, obtained several scale model fan
`acoustic datasets collected from the NASA Glenn 9- by 15-foot Low Speed Wind Tunnel. Of particular importance
`were the datasets collected from scale model representations of the CF6-80E1 fan and the Advanced Ducted
`Propulsor fan.29 The former dataset is significant because the fan was equipped and tested with several stator sets
`that allowed investigations into stator sweep and lean technology. The latter dataset is significant due to its unique
`operation in very low fan pressure ratio regimes. It provided insight into the noise generation mechanisms of these
`types of fans without the masking influence of shock-related sources found in other fans operating in supersonic
`regimes. The fan noise prediction method based on these advanced fan designs is used in this assessment since it is
`more representative of modern, contoured, wide-chord fans for high bypass ratio turbofans. This method was coded
`into ANOPP’s Heidmann fan noise prediction module in 2008, where it now resides as an informal, interim, and
`(currently) undocumented option. (This method is accessed in ANOPP by setting HDNMTH=4.)
`The fourth difference in the noise prediction procedure is the replacement of fixed fan liner acoustic performance
`data with a parametrically varying fan acoustic treatment model. In Ref. 4, the benefits of fan acoustic liners were
`
`modeled by applying an acoustic suppression performance “map” of 1/3rd octave band sound pressure level
`decrements to the predicted hardwall fan source spectra. The liner suppression map was based on measured wind
`tunnel data from the 22-inch diameter “Fan 1” rig in NASA Glenn’s 9- by 15-foot Low-Speed Wind Tunnel.30 The
`most effective treatment tested in these experiments was a double degree of freedom liner applied to the inlet,
`interstage, and aft bypass duct areas. This level of liner suppression was deemed appropriate for use in turbofans of
`this class for the N+1 timeframe. However, this approach did not account for the variability in liner effectiveness
`with changes in inlet and bypass duct dimensions. Inlet diameter and length, and aft bypass exhaust duct height and
`length are important variables in liner performance, and they vary from engine to engine in this turbofan design
`space exploration. ANOPP’s built-in empirical acoustic treatment model31 reacts properly to changes in these
`dimensions, but it underpredicts the level of liner performance that would be expected by the N+1 timeframe.
`Therefore, a hybrid approach is used in this assessment. ANOPP’s built-in treatment prediction method is used since
`it reacts properly to changes in treatment dimensions. But, the inlet and exhaust duct length inputs are deliberately
`exaggerated by ten percent. This artificial increase in duct lengths results in predicted liner spectra that match the
`advanced liner performance levels measured in the rig tests, while retaining the dependency on geometry. In other
`words, the duct length inputs are used as technology calibration factors to force the liner prediction method to match
`the performance of newer, higher performance liners.
`Two modifications were made to the way propagation effects are computed. In 2012, NASA began discussions
`with ICAO’s Noise Technology Independent Expert Panel and contributed to a related task.32 Over the course of the
`discussions, panel members noted that we used a value for the ground flow resistivity (an important parameter in
`ground reflections, calculated in ANOPP via the Chien-Soroka method33) in our previous studies that was at the high
`end of the range typical for grass-covered ground. Regulations in the ICAO Environmental Technical Manual34
`permit the microphones to be located in areas of grass as high as 8 cm. In this study, the value for the flow resistivity
`was reduced from 485 slug/s-ft3 to 291 slug/s-ft3 (the approximate minimum measured for grass-covered ground35)
`reducing the strength of ground reflections. The second propagation effect modification is a credit taken at the
`lateral observer for excess ground attenuation. For observers laterally displaced on a sideline relative to the runway,
`there are additional propagation effects that should be considered. The attenuation is due to differences in ground
`effects (i.e., in surface absorption and reflections), meteorological effects (such as wind and atmospheric gradients),
`and effects due to the airplane configuration (such as engine-airframe shielding and reflections). These effects are
`often collectively known as lateral attenuation or excess ground attenuation. An empirical curve recommended by
`the SAE36 was used to reduce the lateral EPNL to account for these effects. At the lateral sideline distance of 1476
`ft, and where the lateral EPNL is highest, the excess ground attenuation is just over 1 EPNdB.
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`The final noise analysis diflerence is the elimination of advanced airframe noise reduction technologies. In Ref.
`4, reductions in the predicted levels of landing gear, flap, and slat noise were assumed that reflected use of airframe
`technologies such as gear fairings, continuous moldline flexible flap links, and slat cove fillers, respectively. During
`the dialogues with the ICAO Panel mentioned in the preceding paragraph, members advised NASA that these
`advanced airframe noise reduction technologies are too immature to be in use by the N+1 timeframe. Therefore, in
`this study all airframe source noise is computed by ANOPP using the Fink method” without adjustment.
`
`IV. Analysis Results
`
`A. Engine Design
`Overall impacts of the changes to the engine design and modeling assumptions are shown graphically in Figs.
`through 6 for key propulsion characteristics. The results from Ref. 4 are shown as faded lines in the figures. From
`Fig. 3 it is clear that the new engine modeling has reduced engine weight across all twelve of the study engines. The
`decrease is largest for the direct drive engines due to the new LPT design approach used, which limits the number of
`LPT stages and the growth in weight of the direct drive engines at low FPR. Even though the weight penalty for low
`FPR direct drive engines is less now, it is still higher than for the geared engines. (Note that in Ref. 4 the direct
`drive, FPR=l.4 engines were considered impractical designs due to design ground rules resulting in 13 to 15 LPT
`stages, and the results for these designs were shown as dashed lines in the figures. With the current approach, the
`number of LPT stages is limited to nine and these designs are now practical, although heavy. The geared, FPR=l.3
`engine was also previously considered impractical due to integration issues for an under-wing installation. However,
`since the conclusion of the earlier study, a more detailed UHB engine integration trade study was conducted, which
`found that an engine with a nacelle diameter of 9.4 it could be integrated on the ASAT airfiame. Since the FPR=l.3
`case in this study has a nacelle diameter of 9.3 it, this engine is now considered a potential practical design.)
`12000
`
`
`
`
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`Englned-NacelleWeight,lb
`
`10000
`
`3000
`
`6000
`
`4000
`
`2000
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`0
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`1.2
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`N
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`>>-
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`-o-Direct Drive. High
`-I-Geared, High
`_e_D,ed Dm,e_ Low
`1.3
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`1.4
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`1.5
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`(faded ines are Ref. 4 results)
`1.6
`1.7
`1.8
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`Top-of-Climb Fan Pressure Ratio
`
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`DownloadedbyJeffreyBenononAugust19.2013|http://arcaiaaorg|D01:102514/62013-4330
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`Figure 3. Variation of engine+naeelle weight with engine type and design fan pressure ratio.
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`Because the engine design thrust conditions are not the same as in the previous work, a comparison of engine
`thrust-to-weight is more appropriate than absolute weight. In Fig. 4. the thrust-to-weight at the rolling takeoff
`condition (a key engine sizing point) is compared. The improvement in thrust-to-weight ratio for the new designs is
`significant, on the order of 25% for the direct drive engines and 10% for the geared engines.
`Although the new engines are lighter, the thrust spec