9th AIAA Aviation Technology, Integration, and Operations Conference (ATIO) (cid:13)<br>and(cid:13)<br>Air
`21 - 23 September 2009, Hilton Head, South Carolina
`
`AIAA 2009-6942
`
` Analysis of Turbofan Design Options for an Advanced
`Single-Aisle Transport Aircraft
`
`Mark D. Guynn*
`NASA Langley Research Center, Hampton, VA, 23681
`
`Jeffrey J. Berton,† Kenneth L. Fisher,‡ William J. Haller,§ and Michael T. Tong**
`NASA Glenn Research Center, Cleveland, Ohio, 44135
`
`and
`
`Douglas R. Thurman††
`Army Research Lab, Cleveland, Ohio, 44135
`
`The desire for higher engine efficiency has resulted in the evolution of aircraft gas
`turbine engines from turbojets, to low bypass ratio, first generation turbofans, to today’s
`high bypass ratio turbofans. It is possible that future designs will continue this trend, leading
`to very-high or ultra-high bypass ratio (UHB) engines. Although increased bypass ratio has
`clear benefits in terms of propulsion system metrics such as specific fuel consumption, these
`benefits may not translate into aircraft system level benefits due to integration penalties. In
`this study, the design trade space for advanced turbofan engines applied to a single-aisle
`transport (737/A320 class aircraft) is explored. The benefits of increased bypass ratio and
`associated enabling technologies such as geared fan drive are found to depend on the
`primary metrics of interest. For example, bypass ratios at which fuel consumption is
`minimized may not require geared fan technology. However, geared fan drive does enable
`higher bypass ratio designs which result in lower noise. Regardless of the engine architecture
`chosen, the results of this study indicate the potential for the advanced aircraft to realize
`substantial improvements in fuel efficiency, emissions, and noise compared to the current
`vehicles in this size class.
`
`Nomenclature
`
`= Aerodynamic Design Point
`ADP
`ANOPP = Aircraft Noise Prediction Program
`ASAT
`= Advanced Single-Aisle Transport
`BPR
`= Bypass Ratio
`EIS
`= Entry-Into-Service
`EPNL
`= Effective Perceived Noise Level
`FAR
`= Federal Aviation Regulations
`FLOPS = Flight Optimization System
`FPR
`= Fan Pressure Ratio
`HPC
`= High Pressure Compressor
`HPT
`= High Pressure Turbine
`LPC
`= Low Pressure Compressor
`
`
`* Aerospace Engineer, Aeronautics Systems Analysis Branch, MS 442, Senior Member AIAA.
`† Aerospace Engineer, Multidisciplinary Design & Optimization Branch, MS 5-11.
`‡ Aerospace Engineer, Multidisciplinary Design & Optimization Branch, MS 5-11, Member AIAA.
`§ Aerospace Engineer, Multidisciplinary Design & Optimization Branch, MS 5-11, Member AIAA.
`** Aerospace Engineer, Multidisciplinary Design & Optimization Branch, MS 5-11, Member AIAA.
`†† Aerospace Engineer, Multidisciplinary Design & Optimization Branch, MS 5-11.
`
`
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`American Institute of Aeronautics and Astronautics
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`This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States.
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`= Low Pressure Turbine
`LPT
`= Landing-Takeoff Cycle
`LTO
`= Numerical Propulsion System Simulation
`NPSS
`= Operating Empty Weight
`OEW
`= Overall Pressure Ratio
`OPR
`PDCYL = Point Design of Cylindrical-bodied aircraft
`TOC
`= Top-Of-Climb
`TSFC
`= Thrust Specific Fuel Consumption
`UHB
`= Ultra-High Bypass ratio
`WATE
`= Weight Analysis of Turbine Engines
`
`
`A
`
`I. Introduction
`S aircraft manufacturers Boeing and Airbus continue to develop and mature new twin-aisle, wide body aircraft
`designs in the 210-350 seat class, for scheduled first deliveries in 2010 and 2013 respectively, it is anticipated
`that the next major development undertaking for both companies will be a new narrow body aircraft in the Boeing
`737/Airbus A320 class. Boeing and Airbus have been engaged in studies to investigate replacement designs for the
`737 and A320, and published reports indicate that both manufacturers are depending on a next generation engine to
`power these new designs.1 What has yet to be decided is the most attractive advanced engine design for this class of
`aircraft in light of the current metrics of interest in the aviation industry.
`The large fuel consumption and operating cost reductions necessary to make a new single-aisle transport design
`economically viable will require substantial improvements in propulsion system efficiency. In the past, the desire for
`higher engine efficiency has resulted in the evolution of aircraft gas turbine engines from turbojets (bypass ratio
`(BPR) of 0), to low bypass ratio, first generation turbofans (BPR=1-2), to today’s high bypass ratio turbofans
`(BPR=5-10). It is possible that engines for the 737/A320 replacement will continue this trend, leading to very-high
`or ultra-high bypass ratio (UHB) engines. Because of the potential for improved propulsive efficiency, and the
`complementary benefit of lower engine noise, the use of UHB engines has been studied many times over the past
`several decades and there are numerous publications addressing the topic. References 2 through 4 provide a few
`examples. Results published over the years include both positive and negative assessments of UHB engines,
`depending on the assumptions made and the metrics of interest.
`Over time the baseline technologies, market environment (e.g., fuel cost), metrics of interest, and target
`applications change, dictating that concepts such as the UHB engine be periodically revisited. In recent years, fuel
`efficiency, emissions, and noise have become key metrics for aircraft/engine performance. Rising fuel costs have
`greatly elevated the importance of fuel efficiency to the overall profitability of airlines and the success of an aircraft
`design. Noise and emissions are also projected to be of increasing importance in aircraft design as the demand for air
`travel grows. Substantial reductions in aircraft noise and emissions are required to enable unconstrained aviation
`growth without a sharply increasing negative impact on the environment. The 737/A320 class aircraft considered in
`this study represent a significant portion of the global airline fleet. Sixty-five percent of the new aircraft produced
`over the next 20 years are projected to be in this class.5 Advances made to reduce the noise and emissions of these
`aircraft could provide a considerable positive contribution to the goal of minimizing the future environmental impact
`of aviation.
`
`II. Study Objectives and Approach
`The primary objective of this advanced single-aisle transport (ASAT) engine concept study was to determine if
`the thrust specific fuel consumption (TSFC) and noise benefits of higher bypass ratio engines translate into overall
`aircraft system level benefits for a 737 class vehicle entering service in the 2015-2020 time frame. (The scope of this
`study was limited to ducted turbofan engines, open rotor designs may also be viable candidates for a future ASAT
`aircraft and are the focus of a separate study.) The approach taken was to develop a series of analytical engine
`models, apply them to a common airframe model, and assess the overall performance and noise characteristics. The
`main parameter of interest for the study was design fan pressure ratio (FPR). Bypass ratio is inversely proportional
`to fan pressure ratio. As fan pressure ratio is reduced, to maintain thrust fan mass flow must increase, which results
`in higher bypass ratio. It was quickly determined during the initial stages of the study that other key engine design
`choices have a significant impact on the effects of fan pressure ratio and the number of trade parameters was
`expanded. The study was conducted in three analysis “spirals” having different design ground rules and
`assumptions. Each spiral resulted in 16 different configurations for a total of 48 engine/airframe combinations which
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`were analyzed for performance and noise characteristics. This paper presents a brief overview of the study and a
`summary of the results, additional details are provided in reference 6.
`
`III. Modeling and Analysis Methodology
`
`A. Propulsion System Modeling
`Since the propulsion system was the primary area of focus for this study, a substantial amount of effort was
`applied to building analytical models of the study engines. Developing models which were adequately representative
`of engines that could be available for a 737/A320 replacement aircraft was an important objective. However, just as
`important was the requirement of consistency among the engine models. After reviewing available material on
`projected advanced propulsion technologies for the 2015 timeframe, the propulsion systems analysis team developed
`a common design approach and set of technology assumptions which were utilized throughout to enable this
`consistency. The unique characteristics of individual engine architectures may make some assumptions less
`appropriate for certain engines types. This makes applying consistent ground rules and technology assumptions
`across such a wide range of engine designs problematic. The degree to which the resulting study engines are truly
`equivalent in technology and design optimality is uncertain.
`The basic engine architecture for all the engines in this study was a two spool, separate flow turbofan. The
`variations evaluated included the fan drive approach (geared vs. direct drive), the fan pressure ratio, the low spool-
`high spool compression work split, the type of fan nozzle (fixed or variable geometry), the overall pressure ratio,
`and the design Mach number. For a given analysis spiral, all engines were developed with the same Aerodynamic
`Design Point (ADP) (Mach number, altitude, and thrust) and same overall pressure ratio at that point. The ADP was
`selected to represent a nominal top-of-climb (TOC) condition for the advanced airframe. Although for a given spiral
`the overall pressure ratio is the same for all the engines, two different compressor work splits were considered. For a
`given fan pressure ratio and overall pressure ratio, the “low work” engines have a lower pressure rise across the low
`pressure compressor (and a higher pressure rise across the high pressure compressor) compared to the “high work”
`engines. Inlet mass flow for each engine was selected to achieve the net thrust requirement at ADP. In addition to
`meeting a thrust target at TOC conditions, a SLS thrust target of 23,000 lb (hot day, ISA+27°F) was also met by
`adjusting design point burner fuel-to-air ratio. Low fan pressure ratio engines inherently have a greater loss of thrust
`with airspeed (thrust lapse) than high fan pressure ratio engines. To achieve equal ADP thrust capability, the low fan
`pressure ratio engines are operated at higher temperatures. The ADP operating temperatures for the low fan pressure
`ratio engines were below the maximums allowed for the materials assumed, but the higher temperatures could still
`lead to shorter engine hot section life and greater maintenance requirements than the high fan pressure ratio engines.
`Engine life and maintenance issues were not assessed as part of this study. For low fan pressure ratio engines, a
`variable area fan exhaust nozzle was needed to maintain adequate fan surge margin. Throat area of the variable area
`nozzle was varied at off-design to maintain the fan operating conditions equal to, or very close to, the fan peak
`efficiency operating line. Cycle analysis for the engines was performed with the NPSS (Numerical Propulsion
`System Simulation) code.7-9 Analysis of the aeromechanical characteristics and estimates of the engine weight
`(including fan gearbox if applicable) were performed with the WATE (Weight Analysis of Turbine Engines) code.10-
`12 Estimates for engine NOX emission indices (grams of NOX emitted from the engine per kilogram of fuel
`consumed by the engine) were obtained from a correlation developed by NASA combustor technologists during the
`latter stages of NASA’s Ultra-Efficient Engine Technology program.
`
`B. Aircraft Sizing Analysis
`To evaluate and compare aircraft system level performance, the study engines were combined with an advanced
`technology, single-aisle commercial transport airframe model. The aircraft sizing and synthesis computer code
`FLOPS (Flight Optimization System)13 was used as the primary aircraft level sizing and analysis tool. Since the
`objective of the study was a comparison of engine concepts, the primary modeling focus was the propulsion system.
`However, inaccuracies in the airframe model can skew the system level impacts of the engine designs and influence
`the overall conclusions. Special sizing considerations introduced by large diameter, UHB engines were addressed
`through simplifying assumptions and enhancements to the FLOPS analysis. Spreadsheet analyses were used to
`determine landing gear length, engine-out drag, and required vertical tail size so that impacts of large diameter
`engines could be properly captured. Enhancements to basic FLOPS capabilities were also made in the structural
`weight and aerodynamics areas. The wing and fuselage structural weight estimates of FLOPS were replaced with
`estimates from PDCYL. PDCYL offers a less empirical, more analytical weight estimation methodology that is
`more sensitive to parameters such as engine weight and location.14 FLOPS aerodynamic predictions were enhanced
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`through a model calibration process incorporating details of the 737-800 high speed and low speed aerodynamic
`performance.
`1. Baseline Airframe Model
`The Boeing 737-800 (with winglets) was used as a starting point for development of the ASAT airframe model.
`A baseline FLOPS model of a 737-800 like aircraft (162 passenger, mixed-class configuration) was developed using
`a combination of publicly available data on the 737-800 geometry, weight, and performance characteristics;15 a
`CFM56-7B based engine model developed at NASA Glenn; and proprietary aerodynamic data. Model weight
`predictions were calibrated by setting maximum ramp weight and landing weight to the Boeing reported values
`(174,700 lb and 146,300 lb respectively) and comparing the predicted operating empty weight (OEW) to the Boeing
`data. Although the model OEW matched the Boeing data to within 0.5%, calibration adjustments were made to the
`model to match OEW exactly. FLOPS aerodynamic predictions were calibrated to 737-800 high speed aerodynamic
`data. It was not possible to exactly match the 737-800 data at all conditions; however, it was possible to obtain an
`excellent match around the cruise flight conditions. FLOPS predicted mission performance was calibrated to a
`specific point on the 737-800 payload-range diagram provided in reference 15. Prior to adjustments, the FLOPS
`predicted range for the calibration point was ~4% high. Assuming that the mission profile is adequately modeled
`and the aerodynamic model is accurate, the higher FLOPS range is indicative of an under prediction of engine
`TSFC. The NASA-developed engine deck was therefore adjusted so that the FLOPS results matched published
`range capability. Note that it is not possible to separate the impacts of inaccuracies in mission profile, engine TSFC,
`and aircraft L/D when matching range performance. Even though adjustment was only made to the engine model,
`the discrepancy is most likely due to a combination of differences in engine characteristics, aerodynamic
`characteristics, and mission definition. Evaluation and calibration of the FLOPS model was also performed for
`takeoff and landing performance. After some adjustment to the inputs based on 737-800 aerodynamic data, takeoff
`and landing distances for nominal conditions were matched to within ~1.0% of the reported values.
`2. ASAT Airframe Model
`The ASAT airframe model is a derivative of the 737-800 like baseline discussed above, intended to be
`representative of a potential advanced technology replacement aircraft. A conventional airframe-engine layout like
`the 737-800 was assumed based on the hypothesis that unconventional approaches are not sufficiently mature to
`support the expected entry-into-service (EIS) date for this vehicle. The primary airframe technology advancement
`assumed was extensive use of composite materials for the airframe structure. For the Boeing 787 currently in
`development, as much as 50 percent of the primary structure is made of composite materials.‡‡ Other minor
`technology improvements based on the 787 design included an increase in hydraulic pressure and a slight drag
`reduction. Changes were also made to the design mission to reflect performance enhancements projected for an
`advanced aircraft in this vehicle class. Design range (with 32,400 lb payload) was increased from 3060 nm to 3250
`nm. Two cruise Mach numbers were analyzed, 0.72 and 0.80 (typical cruise Mach for the 737-800 is 0.78515). The
`basic 737-800 geometry was not changed for the ASAT model, except for changes in wing sweep corresponding to
`the changes in cruise Mach number.
`3. Propulsion-Airframe Integration
`Propulsion-airframe integration is one of the key considerations for large diameter, UHB engines. Reference 4
`provides an excellent summary of the integration issues associated with large diameter engines and was used as a
`basis for the current study. Concerns highlighted in reference 4 include nacelle drag, ground clearance, windmilling
`drag, thrust reverser operation, and engine placement. These concerns were addressed to varying degrees in the
`study. A simple geometric method was developed to estimate the required landing gear length. Windmilling and
`engine-out drag estimates were made using handbook methods16 and the vertical tail was sized based on
`consideration of both tail volume coefficient and one-engine-out control. Some propulsion-airframe integration
`issues were outside the scope of this study. Examples of issues outside the scope of this study include impacts of
`nacelle diameter on pylon and flap design and potential changes in thrust reverser operation associated with large
`diameter engines. (An estimate of thrust reverser weight was included for all engines, however.)
`
`C. Noise Analysis
`The primary tools used for the noise analysis included: NPSS for the engine cycle analysis; WATE for the
`engine aeromechanical and flowpath analysis; FLOPS for the aircraft trajectory simulation; and ANOPP (Aircraft
`Noise Prediction Program) Level 2617,18 for the source noise prediction and propagation. The NPSS and WATE
`codes were used to generate input data necessary for the ANOPP source noise modeling. Adjustments representing
`noise reduction technologies were made to the source noise spectra prior to propagation. ANOPP noise propagation
`
`
`‡‡ 787 Dreamliner Program Fact Sheet http://www boeing com/commercial/787family/programfacts html Accessed 4/9/2007
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`modeling included spherical spreading, atmospheric attenuation, ground effects, reflections, and lateral attenuation.
`The Effective Perceived Noise Level (EPNL) was calculated at the noise certification points defined in FAR Part
`36.19 EPNL is an integration of the ground observer perceived noise time history which depends on aircraft
`trajectory, noise spectra propagation, frequency integration, and tonal content and amplitude penalties.
`The noise analysis tools were first used to model a 737-800/CFM56-7B and the analytical results were compared
`to noise certification data for that airplane.20 The CFM56-7B was analytically modeled in NPSS using data available
`from several public-domain sources, no proprietary data were used. The thermodynamic, aeromechanical, and
`geometric predictions for the CFM56-7B were used as inputs to ANOPP’s current propulsion source noise
`prediction methods. Good agreement between the certification data and analytical prediction was obtained for the
`lateral (sideline) and approach conditions. Noise at the flyover condition was over predicted by approximately 4
`EPNdB. Through more detailed analysis and comparison of predicted source noise levels to proprietary data it was
`determined that the fan noise predictions might be about 5 dB too high at the flyover, cutback power setting. There
`are many sources of uncertainty in the noise analysis process, however, including the engine cycle and
`aeromechanical modeling (NPSS and WATE), the trajectory and throttle setting assumptions, and numerous other
`potential discrepancies. Because the exact cause of the error cannot be readily determined and the level of error in
`the results was deemed acceptable for this comparative study, no attempt was made to calibrate the noise analysis
`tools and eliminate the discrepancy between predicted and actual 737-800 noise levels.
`A series of advanced noise reduction technologies were applied to the study configurations consistent with the
`2015-2020 EIS target for the vehicle. Chevrons were applied to all core nozzles and to all fixed-area bypass nozzles.
`Chevrons were not applied to bypass nozzles of the low fan pressure ratio engines with variable area nozzles due to
`potential conflict with the variable area nozzle design. Jet noise benefits of the nozzle chevrons were determined
`analytically using the 2004 Stone jet noise prediction method in ANOPP.21 This method is based on 1997 acoustic
`measurements of chevron-equipped nozzles from NASA Glenn’s Aeroacoustic Propulsion Laboratory’s Nozzle
`Acoustic Test Rig freejet facility.22 Conventional inlet, interstage, and aft fan duct liners were applied to reduce fan
`inlet and discharge noise. The benefits of these liners were modeled by applying an acoustic suppression “map” of
`1/3rd octave band sound pressure level decrements to the hardwall fan source spectra predicted by ANOPP. This
`approach differs from the 737-800/CFM56-7B validation study described above, where ANOPP’s built-in treatment
`suppression prediction module23 was used, since a more aggressive treatment configuration would likely be used in
`an advanced engine. The liner suppression map was based on measured acoustic data of 22-inch diameter fan test
`articles in NASA Glenn’s 9×15 Low Speed Wind Tunnel.24 In addition to conventional liners, two advanced
`technologies were applied for fan noise reduction; soft vane stators and over-the-rotor foam metal treatment.25 Both
`of these technologies are applications of acoustic treatment in areas of the engine which currently do not have
`treatment: the fan vanes and above the fan rotor tips. Acoustic tests of both of these technologies were conducted at
`NASA Glenn in 2008. Airframe noise reduction technologies included innovative slat cove designs, flap porous tips,
`and landing gear fairings. These technologies are considered mature enough to be commensurate with the assumed
`EIS timeframe.
`Higher bypass ratio, lower fan pressure ratio engines have inherently higher thrust lapse (i.e., available thrust
`decreases more rapidly with increase in aircraft speed). The impact of higher thrust lapse is manifested in changes in
`climb rates, airspeeds, and throttle settings for takeoff and landing trajectories. Certification noise is impacted by
`these trajectory changes since propulsion noise is a strong function of throttle setting, airframe noise is a strong
`function of airspeed, and altitude and distance from the observer strongly affect noise from all sources. Detailed
`departure and approach trajectories were modeled in FLOPS to enable these engine dependent characteristics to be
`captured in the analysis, which in turn enabled the influence of trajectory on the noise results to be properly
`captured.
`
`IV. Study Trade Space
`
`As mentioned previously, the study was conducted in three separate analysis spirals. The primary differences
`between the three spirals were engine overall pressure ratio at ADP and ADP/cruise Mach number. All of the
`engines in Spiral 1 were designed with an overall pressure ratio (at the top-of-climb ADP) of 32. This overall
`pressure ratio is similar to that of the CFM56 engines that are used on the current Boeing 737 and Airbus A320.
`Current technology large engines can have overall pressure ratios above 40. Although technology advances can lead
`to higher overall pressure ratios, higher pressure ratios lead to smaller compressor blades. There are limits to how
`small a compressor blade can be manufactured and the smaller the blade becomes the less efficient it is due to
`exaggerated blade tip clearance losses and Reynolds number effects. For these reasons it is not possible to simply
`scale down a high overall pressure ratio 80,000 lb thrust engine to a 25,000 lb thrust engine. In Spiral 1, a
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`conventional, fairly conservative design approach was taken which enabled a design overall pressure ratio of 32
`while maintaining blade heights greater than 0.5 inches. Design parameters for the Spiral 1 engines are summarized
`in Table 1.
`
`
`Engine Designation
`S1_Lo_dd_fpr14_VAN
`S1_Lo_dd_fpr15_fixed
`S1_Lo_dd_fpr16_fixed
`S1_Lo_dd_fpr17_fixed
`S1_Lo_g_fpr13_VAN
`S1_Lo_g_fpr14_VAN
`S1_Lo_g_fpr15_fixed
`S1_Lo_g_fpr16_fixed
`S1_Hi_dd_fpr14_VAN
`S1_Hi_dd_fpr15_fixed
`S1_Hi_dd_fpr16_fixed
`S1_Hi_dd_fpr17_fixed
`S1_Hi_g_fpr13_VAN
`S1_Hi_g_fpr14_VAN
`S1_Hi_g_fpr15_fixed
`S1_Hi_g_fpr16_fixed
`
`
`Table 1. Spiral 1 Engine Design Parameters
`Fan Drive
`Fan Nozzle
`ADP
`FPR OPR LPC PR HPC PR
`Direct
`Variable
`M0.80/35kft
`1.4
`32
`1.69
`13.5
`Direct
`Fixed
`M0.80/35kft
`1.5
`32
`1.58
`13.5
`Direct
`Fixed
`M0.80/35kft
`1.6
`32
`1.48
`13.5
`Direct
`Fixed
`M0.80/35kft
`1.7
`32
`1.39
`13.5
`Geared
`Variable
`M0.80/35kft
`1.3
`32
`1.82
`13.5
`Geared
`Variable
`M0.80/35kft
`1.4
`32
`1.69
`13.5
`Geared
`Fixed
`M0.80/35kft
`1.5
`32
`1.58
`13.5
`Geared
`Fixed
`M0.80/35kft
`1.6
`32
`1.48
`13.5
`Direct
`Variable
`M0.80/35kft
`1.4
`32
`2.29
`10.0
`Direct
`Fixed
`M0.80/35kft
`1.5
`32
`2.13
`10.0
`Direct
`Fixed
`M0.80/35kft
`1.6
`32
`2.00
`10.0
`Direct
`Fixed
`M0.80/35kft
`1.7
`32
`1.88
`10.0
`Geared
`Variable
`M0.80/35kft
`1.3
`32
`2.46
`10.0
`Geared
`Variable
`M0.80/35kft
`1.4
`32
`2.29
`10.0
`Geared
`Fixed
`M0.80/35kft
`1.5
`32
`2.13
`10.0
`Geared
`Fixed
`M0.80/35kft
`1.6
`32
`2.00
`10.0
`
`Based on projections of a higher overall pressure ratio for an advanced engine of this class, a second set of analyses,
`Spiral 2, was conducted with a more aggressive design approach to enable an increase in overall pressure ratio to 42.
`In particular, the minimum blade height constraint was relaxed. The engine design parameters for Spiral 2 are shown
`in Table 2. Changes in the low pressure compressor (LPC) and high pressure compressor (HPC) pressure ratios
`associated with
`the higher overall pressure
`ratio can be
`seen by comparison
`to Table 1.
`
`Engine Designation
`S2_Lo_dd_fpr14_VAN
`S2_Lo_dd_fpr15_fixed
`S2_Lo_dd_fpr16_fixed
`S2_Lo_dd_fpr17_fixed
`S2_Lo_g_fpr13_VAN
`S2_Lo_g_fpr14_VAN
`S2_Lo_g_fpr15_fixed
`S2_Lo_g_fpr16_fixed
`S2_Hi_dd_fpr14_VAN
`S2_Hi_dd_fpr15_fixed
`S2_Hi_dd_fpr16_fixed
`S2_Hi_dd_fpr17_fixed
`S2_Hi_g_fpr13_VAN
`S2_Hi_g_fpr14_VAN
`S2_Hi_g_fpr15_fixed
`S2_Hi_g_fpr16_fixed
`
`
`Table 2. Spiral 2 Engine Design Parameters
`Fan Drive
`Fan Nozzle
`ADP
`FPR OPR LPC PR HPC PR
`Direct
`Variable
`M0.80/35kft
`1.4
`42
`1.69
`17.7
`Direct
`Fixed
`M0.80/35kft
`1.5
`42
`1.58
`17.7
`Direct
`Fixed
`M0.80/35kft
`1.6
`42
`1.48
`17.7
`Direct
`Fixed
`M0.80/35kft
`1.7
`42
`1.39
`17.7
`Geared
`Variable
`M0.80/35kft
`1.3
`42
`1.82
`17.7
`Geared
`Variable
`M0.80/35kft
`1.4
`42
`1.69
`17.7
`Geared
`Fixed
`M0.80/35kft
`1.5
`42
`1.58
`17.7
`Geared
`Fixed
`M0.80/35kft
`1.6
`42
`1.48
`17.7
`Direct
`Variable
`M0.80/35kft
`1.4
`42
`2.50
`12.0
`Direct
`Fixed
`M0.80/35kft
`1.5
`42
`2.33
`12.0
`Direct
`Fixed
`M0.80/35kft
`1.6
`42
`2.19
`12.0
`Direct
`Fixed
`M0.80/35kft
`1.7
`42
`2.06
`12.0
`Geared
`Variable
`M0.80/35kft
`1.3
`42
`2.69
`12.0
`Geared
`Variable
`M0.80/35kft
`1.4
`42
`2.50
`12.0
`Geared
`Fixed
`M0.80/35kft
`1.5
`42
`2.33
`12.0
`Geared
`Fixed
`M0.80/35kft
`1.6
`42
`2.19
`12.0
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`The design cruise Mach number selected for the Spiral 1 and Spiral 2 advanced vehicle designs was 0.80, compared
`to a long range cruise Mach number of 0.785 for the 737-800. Some have suggested that to increase fuel efficiency
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`American Institute of Aeronautics and Astronautics
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`the replacement aircraft for the Boeing 737 and Airbus A320 families will actually be designed to fly significantly
`slower. Because of environmental and economic pressures, airlines may be willing to give up something in
`productivity (speed) for reduced fuel consumption. For Spiral 3, the cruise Mach number was reduced to 0.72 to
`assess the impact of cruise Mach number on the relative system level performance for the different engine types. (A
`table of the Spiral 3 engine design parameters is not shown since the only difference compared to Table 2 is an ADP
`Mach number of 0.72 instead of 0.80.)
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`V. Summary of Results
`Figures 1 through 5 attempt to consolidate and summarize the large amount of data generated during this study.
`In these figures, the four different basic engine architectures (low work-geared, low work-direct drive, high work-
`geared, and high work-direct drive) have been collapsed into a single curve by plotting the minimum value among
`the four for a given fan pressure ratio and Spiral. The engine configuration to which this value corresponds is also
`indicated. (Note that in some cases the minimum value is not significantly less than that obtained from the other
`engine architectures.) Figure 1, ramp weight, illustrates the weight penalty associated with low fan pressure ratio,
`which was found consistently across the analysis Spirals. Also, there is a clear preference for high work, geared
`designs at fan pressure ratios up to 1.5 and low work, direct drive engines at higher fan pressure ratios. Higher
`overall pressure ratio (Spirals 2 & 3) and lower Mach (Spiral 3) both reduce ramp weight. For block fuel, shown in
`Fig. 2, there is again consistently a penalty for very low fan pressure ratio engines. The reduction in takeoff
`performance for these engines (due to thrust lapse) coupled with higher aircraft weight leads to higher required
`thrust than the nominal 23,000 lb SLS design value. Although engine scaling laws were used to provide approximate
`characteristics for the higher thrust engines required, the low fan pressure ratio cases could potentially benefit from
`redesigned engines that meet takeoff thrust requirements without scaling. The minimum block fuel consumption
`consistently occurs in the 1.55 to 1.6 fan pressure ratio range (analysis was only conducted at 1.5 and 1.6; the
`minimum shown between those two points is the result of curve fitting the data and may not be the exact minimum).
`As with ramp weight, geared engines are preferred below a fan pressure ratio of 1.5 and direct drive engines above.
`Comparing the Spirals it is evident that both higher overall pressure ratio and lower cruise Mach reduce fuel
`consumption. In the block NOX chart, Fig. 3, all the minimum points are high work engines since the low work
`engines have slightly higher NOX emissions. Similar to ramp weight, the trend is for block NOX to decrease with
`increasing fan pressure ratio, at least up to the highest fan pressure ratio analyzed. In the case of block NOX, gearing
`is beneficial up to fan pressure ratio of 1.6. The increase in overall pressure ratio for Spiral 2 significantly increases
`the block NOX, while the lower cruise Mach in Spiral 3 results in a reduction in block NOX. The trends of landing-
`takeoff cycle (LTO) NOX, shown in Fig. 4, are not as consistent as the other metrics. High fan pressure ratio
`certainly leads to higher LTO NOX, but between FPR=1.3 and 1.5 the variation with fan pressure ratio is not
`consistent. LTO NOX results depend on a combination of the engine characteristics and the aircraft sizing results
`(e.g., trade between engine thrust and wing area necessary to meet takeoff performance); therefore, they exhibit
`more variability. For certification noise the dominant factor is clearly fan pressure ratio as evident in Fig. 5. (Noise
`results are presented in terms of the sum of the noise levels at the three certification points, so-called “cumulative
`noise.”) Although the minimu