Jeffrey J. Berton and Edmane Envia

Glenn Research Center, Cleveland, Ohio

Casey L. Burley

Langley Research Center, Hampton, Virginia

An Analytical Assessment of NASA’s N+1 

Subsonic Fixed Wing Project Noise Goal

NASA/TM—2010-216085

February 2010

AIAA–2009–3144
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Jeffrey J. Berton and Edmane Envia

Glenn Research Center, Cleveland, Ohio

Casey L. Burley

Langley Research Center, Hampton, Virginia

An Analytical Assessment of NASA’s N+1 

Subsonic Fixed Wing Project Noise Goal

NASA/TM—2010-216085

February 2010

AIAA–2009–3144

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

15th AIAA/CEAS Aeronautics Conference (30th AIAA Aeronautics Conference)

sponsored by the American Institute of Aeronautics and Astronautics and 

the Confederation of European Aerospace Societies

Miami, Florida, May 11–13, 2009
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This work was sponsored by the Fundamental Aeronautics Program 

at the NASA Glenn Research Center.

Level of Review: This material has been technically reviewed by technical management. 

This report is a formal draft or working 

paper, intended to solicit comments and 

ideas from a technical peer group.

This report contains preliminary fi ndings, 

subject to revision as analysis proceeds.
NASA/TM—2010-216085 1 

An Analytical Assessment of NASA’s N+1 

Subsonic Fixed Wing Project Noise Goal 

Jeffrey J. Berton and Edmane Envia 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Casey L. Burley 

National Aeronautics and Space Administration 

Langley Research Center 

Hampton, Virginia 23681–2199 

Abstract 

The Subsonic Fixed Wing Project of NASA’s Fundamental Aeronautics Program has adopted a noise 

reduction goal for new, subsonic, single-aisle, civil aircraft expected to replace current 737 and A320 

airplanes. These so-called “N+1” aircraft—designated in NASA vernacular as such since they will follow 

the current, in-service, “N” airplanes—are hoped to achieve certification noise goal levels of 32 

cumulative EPNdB under current Stage 4 noise regulations. A notional, N+1, single-aisle, twinjet 

transport with ultrahigh bypass ratio turbofan engines is analyzed in this study using NASA software and 

methods. Several advanced noise-reduction technologies are empirically applied to the propulsion system 

and airframe. Certification noise levels are predicted and compared with the NASA goal. 

Nomenclature 

AFE altitude above field elevation 

EPNL effective perceived noise level 

FAR federal aviation regulations 

ISA international standard atmosphere 

LSWT NASA Glenn Research Center Low Speed Wind Tunnel 

OASPL overall sound pressure level metric 

PNL perceived noise level metric 

PNLT perceived noise level with tone weighting metric 

SLS sea level static 

UHB ultrahigh bypass 

Introduction 

NASA sets aggressive, strategic, civil aircraft noise reduction goals to improve the quality of life for 

millions of people exposed to airport noise. NASA plays a critical role in reducing community noise 

exposure via technology research and by providing an ambitious noise reduction roadmap for U.S. 

aerospace industry. Under NASA’s Fundamental Aeronautics Program, the Subsonic Fixed Wing Project 

has adopted a community noise reduction goal for new, subsonic, single-aisle, civil aircraft expected to 

replace current 737 and A320 airplanes. These so-called “N+1” aircraft—designated in NASA vernacular 

as such since they will follow the current, in-service, “N” airplanes—are hoped to achieve certification 

noise goal levels of 32 cumulative EPNdB under current Stage 4 noise limits (Ref. 1). In noise 

certification parlance, the cumulative, or algebraic, sum of the three certification EPNLs is often used to 

capture the range of operating conditions. NASA’s N+1 noise reduction goal is of national importance, 
NASA/TM—2010-216085 2 

having been accepted by the Federal Aviation Administration for their Continuous Low Emissions, 

Energy, and Noise (CLEEN) initiative. 

NASA has set N+1 goals for other aircraft system performance metrics as well. Relative to B737-

800/CFM56-7B performance levels, goals are set for 33 percent reductions in block fuel burn and takeoff 

field length. Additionally, a goal for oxides of nitrogen reduction is set to 60 percent below the landing 

and takeoff emission stringencies set in 2004 by the Committee on Aviation Environmental Protection 

(CAEP/6). Since these performance metrics are aggressive, contrasting, and often conflicting, achieving 

goal levels for the noise, emissions, fuel burn, and field length metrics simultaneously may not be 

possible. These goals therefore represent four distinct “corners” of the airplane design trade space. A 

balanced, profitable, business-case airplane design may satisfy one or more of these goals, but is unlikely 

to meet the goal of every metric at once. Even more challenging goals are set for future N+2 and N+3 

aircraft. 

The -32 cumulative EPNdB N+1 goal was set in 2005 based on an earlier, unpublished, internal 

NASA study using the best methods and information available at the time. This paper documents a more 

rigorous investigation conducted in 2008 for the Subsonic Fixed Wing Project using the latest technology 

data and higher-fidelity analytical tools. It is an in-depth, self-consistent, systems analysis study of an 

engine and airplane system using accepted NASA tools and methods. The intent of this study is to verify, 

document, and determine progress towards the -32 EPNdB N+1 goal. This paper describes the analysis of 

a notional airplane located in the “low-noise corner” of the design trade space. The airplane design 

showcases an inherently-quiet ultrahigh bypass ratio (UHB) propulsion system combined with innovative, 

advanced, noise reduction technologies that perhaps would not be used in a business-case design. This 

report is a formal documentation of three presentations made at NASA programmatic meetings in 2008 

(Refs. 2, 3, and 4).

This assessment leverages results from a separate NASA system concept study conducted in 2008, in 

which UHB turbofan engines were examined for a notional, N+1, single-aisle transport (Ref. 5). The 

objective of the NASA UHB engine concept study was to determine if the fuel consumption and noise 

benefits of engines having lower fan pressure ratios (and correspondingly higher bypass ratios) translate 

into overall aircraft system-level benefits for a 737 class vehicle. This internal study provided 

independent guidance to NASA program management regarding the new PurePower geared turbofans 

under development by Pratt & Whitney (Ref. 6). 

In Reference 5, the propulsion system design trade space was examined by designing a representative 

family of 48 N+1 UHB engines, analytically installing them on a common airframe model, and 

performing aircraft mission performance and sizing analyses. The independent propulsion design 

parameters investigated were aerodynamic design point fan pressure ratio, overall pressure ratio, fan drive 

architecture (i.e., direct- or gear-driven), bypass nozzle architecture (i.e., fixed- or variable-geometry), 

booster compression work split, and cruise Mach number. Engine and airframe technology projections 

commensurate with a 2015 entry-into-service date were assumed. Aircraft performance characteristics, 

including properties of the NASA project goal metrics (fuel burn, emissions, field length, and noise) were 

analytically computed for each vehicle and are reported in Reference 5. 

An interesting and important aspect of the aircraft-engine system not always considered in noise 

certification predictions is the influence of airplane trajectory and engine throttling on noise. UHB 

turbofans have significantly different thrust lapse characteristics than other turbofans having higher 

specific thrust, resulting in takeoff and approach trajectories and throttle settings that must be modeled 

properly to correctly compute noise. Detailed takeoff and approach trajectory calculations were made for 

each of the airplanes in Ref. 5. The same trajectory modeling procedures are used to compute certification 

noise in this study. 

One of the quietest propulsion systems investigated in Reference 5—a geared turbofan with an 

ultrahigh bypass ratio—is selected for further, refined study and is the basis of this report. It is important 

to note that this UHB engine is not the best performing propulsion system examined in Reference 5. That 

is, of the 48 engine designs studied, it does not result in an airplane having minimum block fuel or takeoff 

gross weight. Its low fan pressure ratio led to a relatively high engine weight and a large nacelle diameter, 
NASA/TM—2010-216085 3 

resulting in propulsion-airframe integration penalties and poorer overall mission performance. Instead, it 

is selected in this study because it represents the “low-noise corner” of the design trade space, noting once 

again that meeting all of the NASA goals simultaneously may not be possible. 

Method of Analysis 

The low-noise propulsion system singled out from Reference 5 for analysis of the noise goal is 

comparable in thrust class to the CFM56-7B engines now in use on the 737. However, it is a more 

advanced, geared, UHB turbofan with a lower design fan pressure ratio and lower fan tip speeds than the 

CFM56. The engine also has a higher overall pressure ratio, better component performance levels, higher 

hot-section temperatures, lower exhaust velocities, and a higher booster compression split than the 

CFM56. Once expected to enter service as early as 2015, it appears the 737 replacement aircraft will be 

delayed for several more years (Ref. 7). Nonetheless, engine component and subsystem performance, 

cooling levels, and material technologies appropriate for an approximate 2015 service entry date are 

assumed. Key computed characteristics of the propulsion system at sea level static and standard day 

conditions are shown in Table 1. The overall pressure ratio is aggressively high (42 at Mach 0.8/35,000 ft, 

and 32 at sea level static conditions) for an engine of this thrust class and comparatively small core. With 

the low-pressure fan spinning slowly at subsonic tip speeds (only 990 ft/s), a fan-drive gearbox is 

necessary to avoid low-pressure turbine design complications. Such low-speed, low-pressure fans may 

have relatively large diameters, presenting potential integration issues. Another consequence of the fan’s 

low pressure ratio requires the engine cycle to have some means of variable geometry in the bypass 

stream to maintain proper fan operability and safety margins. A variable-area bypass nozzle is used to 

maintain a constant surge margin throughout the operating envelope. The thermodynamic engine cycle 

performance is analyzed using the Numerical Propulsion System Simulator code (NPSS (Refs. 8 and 9)). 

NPSS is a variable-fidelity, object-oriented, engine cycle analysis tool developed jointly by NASA and 

U.S. industry. It is currently the accepted, state-of-the-art software for airbreathing engine cycle 

performance analysis for U.S. aerospace industry, academia, and NASA. A multiple design point analysis 

is performed on the engine cycle in order to meet several performance requirements such as airplane 

thrust demand at rolling takeoff and top-of-climb conditions, as well as to set flow rates, cycle 

temperatures, pressures, spool speeds, and cooling levels. The bypass ratio is a fallout of the fan pressure 

ratio and pressure levels in the bypass and core exhaust ducts at the cycle’s aerodynamic design point. 

Aeromechanical design, flowpath, and engine weight analyses are performed with the Weight Analysis of 

Turbine Engines code (WATE (Ref. 10)). WATE has been significantly upgraded since its initial 

introduction in the 1970s and is currently implemented as a suite of NPSS interpretive analysis elements. 

At NASA, WATE is coupled with NPSS to provide a complete modeling capability of turbofan engines. 

Additional details on the engine design may be found in Reference 5. 

TABLE 1.—KEY PROPULSION SYSTEM CHARACTERISTICS 

(SLS CONDITIONS, MAXIMUM THROTTLE, ISA+0). 

Thrust ............................................................................ 23000 lb 

Airflow ......................................................................... 1060 lb/s 

Fan pressure ratio ................................................................. 1.30 

Fan diameter ....................................................................... 82 in 

Fan tip speed ................................................................... 990 ft/s 

Bypass ratio ............................................................................ 16 

Overall pressure ratio .............................................................. 32 

The notional N+1 airplane is based on an analytical evolution of the currently-in-service Boeing 737-

800 with winglets. In Reference 5, a reference 737 analytical model was developed based on publiclyavailable 737-800 geometry, weight, and performance information; proprietary low-speed and clean-
NASA/TM—2010-216085 4 

configuration aerodynamic data; and a NASA NPSS representation of the CFM56-7B engine. The 

CFM56-7B was analytically modeled in NPSS using data available from several public-domain sources, 

such as FAA type certification data sheets, manufacturer-provided operating documents, technical reports, 

Jane’s Aero Engines, and manufacturer’s websites. No company-proprietary data were used. The 

reference 737 carries 162 passengers in a single-aisle, two-class seating arrangement. The aircraft 

synthesis was carried out using NASA’s Flight Optimization System (FLOPS (Ref. 11)) computer 

program. Minor calibrations to the FLOPS-computed component weights and aerodynamics were 

performed to match published operating empty weight and range capability of 3060 nm at a 32,400 lb 

payload. The FLOPS model was set up to perform a basic wing-engine sizing analysis. Scaling the wing 

and engine sizes of the reference model for minimum takeoff gross weight subject to active aircraft 

performance constraints led to results that were consistent with the actual 737 aircraft. The 737 model 

was further developed to include detailed, low-speed takeoff and landing assessments using FLOPS’s 

built-in, time-stepping trajectory analysis module. Compliance with the airworthiness requirements 

described in Part 36 and 25 of the Federal Aviation Regulations (Refs. 1 and 12, respectively) is observed. 

Takeoff and landing performance data (Ref. 13) for standard day, dry runway conditions were used for 

validation of the FLOPS low-speed trajectory model. Takeoff and landing distances matched to within 

approximately one percent of the reported values. 

To transform the reference 737 into the advanced N+1 vehicle model, the design cruise Mach number 

is increased from 0.785 to 0.800, with an appropriate increase in wing sweep to reflect the higher 

airspeed. The wing aspect ratio and taper ratio are unchanged. The 162-passenger, mixed-class, singleaisle cabin arrangement is maintained, but the design range at the 32,400 lb payload point is increased 

from 3060 nm to 3250 nm. The performance improvements in airspeed and range are considered 

appropriate for a future vehicle in this class. Broad use of composite structural materials is assumed 

relative to the all-metal construction of the 737. Composite construction of primary structures is assumed 

to result in a 15 percent reduction in the component weights computed for the wing, fuselage, and 

empennage. This is comparable to the structure of the new Boeing 787, where as much as 50 percent of 

the primary structure is made of composites (Ref. 14). Additional technology improvements similar to 

those found on the 787 include an increase in hydraulic pressure to 5000 psi and a 1 percent reduction in 

drag due to trailing edge variable camber and drag clean-up. A Vehicle Sketch Pad (Ref. 15) planform 

schematic of the N+1 vehicle (without engines) is shown to scale in Figure 1. 

 Figure 1.—Planform view of the N+1 airplane. 
NASA/TM—2010-216085 5 

Propulsion-airframe integration effects are important to address properly for large-diameter UHB 

engine nacelles. Low-speed, low-pressure fans—such as the 82-in. diameter fan considered here—present 

unique propulsion-airframe integration challenges. Nacelle drag, ground clearance, windmilling drag, 

vertical tail sizing, thrust reverser operation, engine placement, and gear length and weight issues are 

considered. The methods used to capture these effects are described in greater detail in Reference 5. Like 

the reference 737 model, the N+1 vehicle model is parametrically sized for minimum takeoff gross 

weight, and detailed takeoff and landing trajectories are computed. 

The noise analysis approach, methods, assumptions, and tools used in this study have been examined 

by acoustics experts at NASA Glenn, NASA Langley, U.S. industry, and academia as part of a 

comprehensive, multifidelity, NASA acoustic tool benchmarking activity (Ref. 16). Efforts to benchmark 

the accuracy of NASA’s aircraft-system-level noise prediction code for airport community noise are also 

leveraged (Ref. 17). This computer code—the Aircraft Noise Prediction Program (ANOPP, Release Level 

26 (Refs. 18 and 19)—is a systems-level code used in this study to compute certification noise for the 

N+1 vehicle. The certification noise predictions of the reference 737 described in Reference 17 serve as a 

validation of the methods and tools used in this N+1 vehicle study. 

Freefield, lossless, 1/3rd octave band frequency component source noise levels are computed using 

predictive modules within ANOPP. The UHB engine’s thermodynamic, aeromechanical, and geometry 

data are used as inputs to ANOPP’s propulsion source noise prediction methods. At NASA, the NPSS and 

WATE airbreathing component element libraries have functions, viewers, and case files coded in 

interpretive language to produce engine data to be used as ANOPP input parameters. These so-called 

“engine state tables” are the preferred method to transfer engine state data to ANOPP’s source noise 

prediction modules. Engine state data—consisting of pressure, temperature, flow area, spool speed, and 

fuel and air flow rates—are computed by NPSS for a range of airspeeds, altitudes, and throttle settings at 

standard acoustic day (ISA+18 °F) conditions. As the airplane traverses its flight path, engine data at the 

appropriate airspeed, altitude, atmospheric conditions, and throttle setting are interpolated from the state 

tables and are delivered to the source noise prediction modules. 

The UHB engine’s jet noise sources are predicted using the Stone method (Ref. 20). This method—

implemented in ANOPP’s “ST2JET” module—is perhaps the best semi-empirical jet noise prediction 

method developed to date. It uses an innovative, physics-based, Bayesian regression of jet acoustic data 

that seamlessly spans various nozzle flow and geometry regimes. It is calibrated against jet acoustic data 

collected from several types of axisymmetric nozzle designs, such as single-flow and coannular nozzles, 

plug and plugless nozzles, and convergent and convergent-divergent nozzles. The method is considered in 

the acoustic tool benchmarking activity described above to be accurate (Ref. 21). Coannular nozzles with 

bypass ratios of up to 15 are part of Stone’s calibration, and is therefore considered quite applicable to 

UHB engines. The method accepts nozzle flow characteristics as input parameters from the engine state 

tables described above. The Stone method also captures the noise-reduction effects of nozzle perimetermixing devices, such as chevrons. This feature of the 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 (Ref. 22). The Stone perimeter-mixing method predicts reductions in large- and 

intermediate-scale jet mixing noise and an increase in small-scale (high frequency) jet mixing noise near 

the nozzle exit plane. Some inaccuracies of the Stone perimeter-mixing method for certain nozzle designs 

to which the method was not calibrated are noted in Reference 21. However, since the N+1 UHB nozzle 

is coannular with an external plug, and is much the same as the chevron nozzles used in Stone’s 

regression calibration (i.e., the “3IB” and “3IC” coannular nozzle configurations of Ref. 22), the method 

is considered accurate for use here. Chevrons are assumed present for the UHB engine’s central core 

nozzle. However, chevrons are not applied to the bypass nozzle due to potential conflict with the 

actuation system needed for the variable-area bypass nozzle design. The N+1 UHB nozzle is proposed to 

be similar in architecture to the 3IB nozzle test configuration (with a nominal bypass ratio of 5) shown in 

Figure 2, but with a much higher bypass ratio. The 3IB nozzle has 12 in-flip core chevrons and—unlike 

the 3IC nozzle—no outer bypass chevrons. 
NASA/TM—2010-216085 6 

 Figure 2.—3IB chevron coannular nozzle model. Figure 3.—Scale model Advanced Ducted Propulsor 

 “Fan 1” test article installed in the NASA LSWT. 

The UHB engine fan noise source is modeled using experimental acoustic data collected from a scalemodel fan test article in NASA Glenn’s 9 ft by 15 ft Low Speed Wind Tunnel (LSWT).23 The fan tested 

is a 22- in. diameter model of Pratt & Whitney’s Advanced Ducted Propulsor, shown in Figure 3. 

Ordinarily in an aircraft system noise prediction problem, fan source noise would be computed using one 

of ANOPP’s built-in methods. Indeed, in the study upon which our engine is based (Ref. 5), fan noise was 

modeled using ANOPP’s internal General Electric empirical method (Ref. 24), which is considered to be 

generally accurate (Ref. 25). Using a built-in method is convenient in an aircraft systems study, since 

system effects such as fan operation, speed, scale, and convective amplification effects that vary with fan 

design, engine throttle position, and flight condition are properly accounted for by the model without need 

for user intervention. A built-in method is particularly helpful when performing a design space 

exploration of an engine architecture such as in Reference 5, where fan design parameters such as 

pressure ratio are parametrically varied. However, while the method described in Reference 24 may be 

appropriate to predict noise during a design space exploration of many fans, greater accuracy may be 

obtained if the fan to be analyzed is very similar to a fan already tested in an aeroacoustic facility. 

Moreover, the Advanced Ducted Propulsor “Fan 1” tested in the LSWT is representative of modern, 

contoured, wide-chord fans for high-bypass engines, whereas the built-in ANOPP methods are calibrated 

to older, narrow-chord, straight-bladed fans. Fan 1 is an 18-blade, high bypass, scale-model fan with a 

low design pressure ratio of 1.29. This stage pressure ratio compares well with the 1.30 pressure ratio of 

the N+1 UHB fan at similar conditions, and the blade count is also identical. Although the fan loadings 

are different (the UHB fan of Ref. 5 and the scale model Fan 1 SLS corrected tip speeds are 990 and 

840 ft/s, respectively), spool speed in thermodynamic cycle analysis is only an artifice when scaled 

component performance maps are used. To use the noise measured from the Fan 1 test article, the tacit 

assumption is that the N+1 UHB fan is operating at the lower, 840 ft/s, corrected takeoff tip speed. There 

are, of course, aeromechanical, flowpath, and weight implications that cannot be so easily ignored, but the 

engine cycle performance is unchanged. 

The measured noise levels of Fan 1 were collected and reduced to lossless, freefield, 1/3rd octave band 

spectra by methods described in Reference 23. These spectra must be modified further before they may be 

used in an aircraft system noise problem. First, an atmosphere-dependent source strength correction is 

applied to bring the levels from facility conditions to standard atmospheric pressure and temperature. 

Sources other than pure fan noise (such as low-frequency airflow scrubbing and facility noise sources, 

and high-frequency measurement anomalies) are subtracted from every spectrum. The levels are adjusted 

from model-scale, 22-in. fan diameter to full-scale, 82-in. diameter by applying amplitude and frequency 

shifts. Finally, convective amplification Doppler effects are added as the spectra are analytically “flown” 

at variable airspeed, ranging from static to Mach 0.4. The end result is a database of fan noise levels, with 
NASA/TM—2010-216085 7 

discrete independent parameters consisting of 51 emission polar (yaw) angles, 8 shaft speeds, and 5 flight 

Mach numbers. The source is assumed to be symmetric in emission azimuthal (roll) angle. The spectra are 

tabulated for interpolation and fed into the aircraft system noise simulation via ANOPP’s Acoustic Data 

Module; a feature that allows users to create their own noise sources to be used instead of using ANOPP’s 

internal source noise prediction methods. 

The acoustic data collected from Fan 1 are hardwall; i.e., no acoustic treatment is present. 

Conventional inlet, interstage, and aft fan duct liners are applied to reduce fan inlet and discharge noise. 

The benefits of these liners are modeled by applying an acoustic suppression performance “map” of 1/3rd

octave band sound pressure level decrements to the hardwall fan source spectra. This approach differs 

from the 737-800/CFM56-7B validation study described in Reference 17—where ANOPP’s built-in 

treatment suppression prediction module was used (Ref. 26)—since a more aggressive treatment 

configuration would likely be used in an advanced UHB engine. The liner suppression map is based on 

measured wind tunnel data from the 22-in. diameter Fan 1 rig in NASA’s LSWT (Ref. 27). The most 

effective treatment tested proved to be double degree of freedom liners applied to the inlet, interstage, and 

aft bypass duct areas. These liners are tuned to attack the discrete interaction tone’s second harmonic at 

maximum takeoff-rated power. An initial treatment suppression map is generated based on the measured 

differences between the treated and hardwall measurements. This simple map is scaled from the 22-in. 

model dimensions to 82-in. diameter full scale using standard Part 36 regulation frequencies. The data are 

smoothed with a regression technique to fit a modified Weibull probability density function as shown in 

Figure 4. A Weibull-shaped curve is chosen since liner effectiveness is small at low frequencies and 

increases to a maximum near its tuned frequency. Low-frequency liner self noise—caused by air flow and 

surface roughness, and can be seen in the Figure 4—is small and is ignored by this regression. The 

logarithm of the ratio of inlet diameter to wavelength is an appropriate regression choice for the 

independent parameter. The complete treatment suppression map is shown in Figure 5. Maximum 

effectiveness is approximately 12 dB at an emission polar (yaw) angle of 99° from the inlet. 

Further adjustments representing fan noise reduction technologies are made to the hardwall fan source 

noise spectra prior to propagation. In addition to conventional fan liners, two advanced technologies are 

applied: soft vane stators (Ref. 28) and over-the-rotor foam metal treatment (Ref. 29) Both of these 

technologies are applications of acoustic treatment in areas of the engine which currently do not have 

treatment: on the surface of the fan vanes and above the fan rotor tips. Soft vane stators have small 

Helmholtz resonator chambers inside the airfoils and are covered by a porous surface material over a 

portion of the chord near the leading edge. This single-degree-of-freedom treatment reduces the unsteady, 

noise-generating, pressure fluctuations like other turbofan treatment panels, but acts directly on the vanes. 

Figure 4.—Sample regression fit of acoustic liner performance 

(emission yaw angle 95° from inlet). 

-14

-12

-10

-8

-6

-4

-2

0

2

0.0 0.5 1.0 1.5 2.0 2.5

Logarithm of Inlet Diameter to Sound Wavelength Ratio

Treatment Suppression, dB

Suppression Data

Regression Fit
NASA/TM—2010-216085 8 

Over-the-rotor foam metal is a bulk sound-absorbing material situated over the fan rotor where, normally, 

conventional fan rub strips are located. The metal foam is integrated with the fan casing containment 

structure and provides a greater effective treatment area. An illustration of these two technologies is 

shown in Figure. 6. 

These technologies are anticipated to have matured enough for incorporation in the UHB engine 

assumed in this study. These technologies are also assumed to be relatively lightweight, inexpensive, lowmaintenance, and free of aerodynamic performance penalties such that an engine manufacturer would be 

willing to make use of them on their product. The treatment in these locations attenuates both fan 

broadband and interaction tone noise and are anticipated to work together without conflict to produce 

additive fan noise reduction benefits. Acoustic tests of both of these technologies were conducted at 

NASA Glenn in 2008. Based on these tests, a system-level decrement of –4 dB is applied to the freefield 

hardwall fan source noise levels. Until these technologies mature and their impact is better understood, 

this adjustment is applied as a simple constant to the fan sound pressure levels—already reduced by the 

conventional liners—across all 1/3rd octave band frequencies, directions, and throttle settings. 

Figure 6.—Advanced fan noise reduction technologies: over-the-rotor foam metal bulk treatment (left) 

and soft vane stators (right). 

Figure 5.—Overall acoustic liner suppression performance map. 

Treatment

Suppression,

dB

50

150

100

0

-5

-10

Emission Polar (Yaw) 

Angle, deg

0

1

2

Log of Inlet 

Diameter to 

Wavelength 

Ratio
NASA/TM—2010-216085 9 

The remaining propulsion noise sources considered are combustion, turbine, and compressor noise; 

often collectively known as core noise. Historically in noise certification, core noise tends to be 

significant only at the approach certification point. Core noise is revealed at low, approach throttle 

settings when fan and jet noise are reduced due to lower fan tip speeds and lower nozzle exhaust 

velocities. A fundamental turbine discrete interaction tone may also fall under 10 kHz at approach throttle 

and may become problematic. In advanced UHB engines, however, core noise may become significant at 

higher throttle settings as well, since jet and fan noise are lower due to increased bypass ratio, low fan tip 

speeds, and modern, more effective noise reduction technologies. Accurate core noise modeling for UHB 

engines is essential. Core noise is predicted using ANOPP’s built-in procedure (the “GECOR” module) 

developed by Emmerling (Ref. 30). and later modified by Ho (Ref. 31). Preliminary results from NASA’s 

current acoustic tool benchmarking study (Ref. 16) indicate that the ANOPP method agrees well with 

static core noise separated from the overall acoustic signature of the Honeywell TECH977 business-jetclass research engine (Ref. 32). General Electric also evaluated the ANOPP core noise method favorably 

in 1996 based on static acoustic test comparisons to CF6-80C2, QCSEE, and E3 engines, as well as 

comparisons with their own proprietary method (Ref. 24). Like the Stone jet noise method, data are fed 

into ANOPP’s core noise module via engine state table data. Key input parameters are maximum cycle 

pressure, temperature, and flow rate. A shortcoming of the ANOPP core noise method is that the 

maximum overall pressure ratio allowed is 30. This is not unexpected since the method dates to the 1970s 

and was calibrated against engines then having much lower overall pressure ratios. The N+1 UHB engine 

produces a SLS pressure ratio of 32 (already over the ANOPP limit; see Table 1), and increases to 

approximately 34 at an airspeed of 180 ktas; roughly corresponding to the flight condition just before the 

throttle cutback beyond the lateral certification observer. The core noise therefore may be under predicted 

by an unknown amount. No core noise reduction technologies are assumed. 

Like engine core noise, airframe noise sources ordinarily tend to be significant only during approach, 

when engine throttle settings are reduced. With advanced, quiet, UHB engines, however, airframe noise 

sources may be expected to be more prominent at approach, and at higher throttle settings as well, since 

they may become audible over the already-low fan and jet noise sources. In addition, for practical reasons 

and safety considerations, air traffic on approach tends to fly over significant distances at comparatively 

low altitude. This is in sharp contrast to departure trajectories, where the pilot’s intent is to gain altitude 

rather quickly. Remarkably, despite having much louder noise signatures on departure than on approach, 

this behavior results in noise “footprints” that have roughly equivalent enclosed areas on approach and 

departure. These considerations point to the need for aggressive airframe noise reduction technologies for 

UHB-equipped aircraft. Freefield, lossless, 1/3rd octave band spectra for flap, slat, landing gear, and 

trailing edge airframe noise sources are predicted using a method developed in 1977 by M.R. Fink of the 

United Technologies Research Center for the FAA (Ref. 33). The Fink method—programmed into 

ANOPP’s “FNKAFM” module—accepts gross airframe dimensions such as span, flap chord lengths, and 

gear configuration and dimensions; all of which may be obtained from a simple, open-literature, threeview aircraft drawing. Scrutiny of portions of the Fink method to date has not yet revealed any 

fundamental issues (Ref. 34). 

The airframe noise reduction technologies applied to the N+1 airframe are aggressive and 

substantial—fair game, since this study assesses the “low-noise corner” of the aircraft design trade space. 

These technologies are currently considered somewhat immature, but with aggressive funding and 

development, they could become available in time for the N+1 vehicle’s entry into service. The high-lift 

flap arrangement on the N+1 aircraft is assumed to consist of retractable, inboard- and outboardsegmented, slotted flaps. Unlike conventional trailing edge flaps, however, the N+1 flaps are assumed to 

be equipped with continuous moldline links. These links abate the aerodynamic noise generated by 

airflow around the discontinuous flap side edge by extending and continuously fairing a flexible moldline 

into the wing trailing edge. A flow analysis (Ref. 35) illustrating the mechanism and the effect on flap 

side-edge streamlines is shown in Figure 7. The ordinarily strong vortex system is diffused by the 

moldline, thus reducing noise. The noise benefit of continuous moldline links is modeled by subtracting 
NASA/TM—2010-216085 10 

 

8 dB from the conventional flap noise computed by the Fink method. Until this technology matures and 

its impact is better understood, the benefit is assumed constant across all 1/3rd octave band frequencies, 

directions, and airspeeds. 

Another source of aerodynamic noise radiates from the extended leading edge slat cove. Slat noise 

results from complex flow patterns in the slat cove region. If these turbulent, circulating, vortex-shedding 

flow patterns are largely prevented from developing, while still maintaining uniform flow through a slat 

slot, substantial noise reduction benefits may be achieved. An example of a flexible material filling the 

cove of a bench top slat model is shown in Figure 8. The noise benefit of slat cove filler is modeled by 

subtracting 10 dB from the conventional slat noise computed by the Fink method; constant across 

frequency, direction, and airspeed. 

The main gear and nose gear are assumed to have spoilers for flow deflection and fairings to improve 

the aerodynamic shape while still allowing easy access for maintenance and inspection. In addition, 

trailing edge treatment methods are assumed, such as serrated edges or brushes, to reduce trailing edge 

noise. A graph of the assumed benefits of these and the other airframe noise reduction technologies are 

shown in Figure 9. 

Using an assumption of acoustic superposition, the freefield, lossless spectra for all of the noise 

sources described above are analytically summed in the vicinity of the aircraft. Real noise sources are, of 

course, complex, distributed signals that are affected by other acoustic sources, aircraft external surfaces, 

and the environment. No provisions are made to adjust the component spectra for acoustic near-field 

phenomena such as source interactions, reflections, refraction, diffraction, or other effects. 

The summed spectra are propagated to the three certification observers on the ground in accordance 

to the specifications for certification measurements. Noise propagation effects accounted for include 

spherical spreading, Doppler shift and convective amplification, atmospheric attenuation, ground 

reflections based on data for grass-covered ground, and extra ground attenuation. More complex 

propagation phenomena such as scattering, weather effects, and terrain are not modeled. The airplane 

trajectory, computed as described earlier, is fed into the ANOPP simulation. Vector geometry analyses for 

the airplane relative to the three certification microphone measurement locations—shown in Figure 10—

are performed within ANOPP as functions of source time. Note that in the interest of international 

rulemaking terminology harmonization, the former “sideline” certification location term has been 

deprecated in favor of “lateral,” as has “takeoff” to “flyover.” In any event, the propagated acoustic 

spectra are predicted at half-second intervals at each of the three certification locations. From these 

spectra, ANOPP computes OASPL, PNL, PNLT, and other noise metrics of interest as functions of 

observer time. The EPNL certification noise metric is computed from the PNLT-time noise history at 

each observer as prescribed in Reference 1. 

Figure 8.—Example of a pneumatic slat cove filler 

technology concept (bench top model). 

Figure 7.—Flow analysis for continuous moldline link, 

showing surface pressure and side-edge streamlines.
NASA/TM—2010-216085 11 

Figure 10.—Noise certification observer arrangement relative to hypothetical combined takeoff and landing 

aircraft trajectories. 

Results and Discussion 

As noted, the UHB engine selected from those studied in Ref. 5 for refined study here represents the 

“low-noise corner” of the airplane design trade space. As such, it has a very low design fan pressure ratio, 

exhaust velocities, and it has quite different specific thrust characteristics than the turbofans used on 

contemporary 737s. The CFM56-7B series turbofan, currently used on the 737-600, -700, -800, and -900 

models, has a runway ground roll net thrust lapse (i.e., from zero to 170 ktas at sea level) of about only 

20 percent. The UHB engine analyzed here has a runway thrust lapse of nearly 30 percent. Note that not 

all UHB engines would lapse this much in thrust; it may be that a UHB engine design enabling a better 

business-case airplane design might not lapse much more than the CFM56-7B. However, our low-noise 

UHB engine behaves significantly different, resulting in takeoff and approach trajectories and throttle 

settings that must be modeled properly to correctly compute certification noise. This is in contrast to 

many other certification noise estimates in the conceptual phase of aircraft design, where fixed 

trajectories and throttle settings are often assumed based on previously-collected data from other 

representative aircraft. Trajectory data computed for the N+1 airplane and the 737-800 for altitude, 

airspeed, and throttle setting are shown in Figures 11, 12, and 13, respectively. The trajectories shown 

2000 m 

(6562 ft) 

Flyover (with cutback) 

reference 

Lateral 

reference 

Approach 

reference 

6500 m 

(21 325 ft) 

450 m 

(1476 ft) 

Figure 9.—Airframe noise reduction technology benefits. 

-12

-10

-8

-6

-4

-2

0

2

0.1 1 10

1/3 Octave Band Center Frequency, kHz

Technology Suppression, dB

Trailing edge treatments

Main gear spoiler, fairing

Nose gear spoiler, fairing

Continuous moldline links

Slat cove filler
NASA/TM—2010-216085 12 

 

are presented as “analytical touch-and-go” operations with both takeoff and landing data shown 

simultaneously. For presentation purposes, the touchdown point on landing is coincident with the 

point of brake release on takeoff. Calculations are made for a sea level runway at standard acoustic day 

(ISA+18 °F) conditions. The airframe noise reduction technologies noted above are assumed to have no 

impact on the low-speed aerodynamic performance of the airplane. The triangular markers on each plot 

denote the noise certification measurement locations. 

The approach microphone markers are shown in the figures at 6562 ft (2000 m) behind the runway 

threshold (i.e., behind the location of the 50 ft landing obstacle), and 7518 ft from the touchdown point on 

the runway centerline. Note that in Figure 13, the UHB engine throttle setting must be set higher than the 

CFM56-7B (in percent thrust) in order to maintain a 3° approach glide slope. 

The lateral microphone markers are shown in the figures at 14,000 ft from brake release, on a lateral 

sideline displacement distance of 1476 ft (450 m) from the runway centerline. The 14,000 ft location 

corresponds to the distance from brake release where the airplane has reached 984 ft (300 m) AFE, as 

permitted—when approved—by regulation B36.3.a.1 (Ref. 1). A check of other lateral locations less than 

and greater than 14,000 ft from brake release confirms that peak lateral EPNL does in fact occur 

approximately at the 14,000 ft location. Note that in Figure 12, the airspeed of the N+1 airplane is 12 ktas 

slower than the 737 at that point, leading to a longer dwell time for the N+1 airplane and causing the 

duration component of the lateral EPNL to increase somewhat. 

Figure 13.—Throttle setting: 737 and N+1 trajectory 

comparison. 

0.0

0.2

0.4

0.6

0.8

1.0

1.2

-15 -10 -5 0 5 10 15 20 25 30

Distance from Brake Release, 1000 ft

Thrust / Maximum Thrust

B737-800/CFM56-7B26

N+1 Airplane

0.0

0.2

0.4

0.6

0.8

1.0

1.2

-15 -10 -5 0 5 10 15 20 25 30

Distance from Brake Release, 1000 ft

Thrust / Maximum Thrust

B737-800/CFM56-7B26

N+1 Airplane

Figure 12.—Airspeed: 737 and N+1 trajectory 

comparison. 

0

50

100

150

200

250

-15 -10 -5 0 5 10 15 20 25 30

Distance from Brake Release, 1000 ft

True Airspeed, ktas

B737-800/CFM56-7B26

N+1 Airplane

0

50

100

150

200

250

-15 -10 -5 0 5 10 15 20 25 30

Distance from Brake Release, 1000 ft

True Airspeed, ktas

B737-800/CFM56-7B26

N+1 Airplane

Figure 11.—Altitude: 737 and N+1 trajectory 

comparison. 

0.0

0.5

1.0

1.5

2.0

2.5

-15 -10 -5 0 5 10 15 20 25 30

Distance from Brake Release, 1000 ft

Altitude, 1000 ft AFE

B737-800/CFM56-7B26

N+1 Airplane

0.0

0.5

1.0

1.5

2.0

2.5

-15 -10 -5 0 5 10 15 20 25 30

Distance from Brake Release, 1000 ft

Altitude, 1000 ft AFE

B737-800/CFM56-7B26

N+1 Airplane
NASA/TM—2010-216085 13 

The flyover microphone markers are shown in the figures at 21,325 ft (6500 m) from brake release on 

the runway centerline. A noise abatement throttle cutback is used. The engine climb thrust at this point is 

reduced to the minimum level permitted by regulation (i.e., Ref. 1 requires a minimum climb gradient of 

four percent with both engines operating, or level flight with one engine inoperative.). The throttle 

cutback takes place between 16,000 and 17,000 ft from brake release. Note the power cutback takes place 

at approximately 1200 ft AFE. This is above the minimum altitude permitted (i.e., 984 ft, or 300 m AFE 

for a twinjet), in an attempt to gain additional altitude and minimize the noise at the flyover observer. 

Note that the UHB engine throttle cutback is shallower than the CFM56-7B (in percent thrust) in 

order to maintain the necessary minimum cutback climb gradient. The altitude that the N+1 airplane is 

able to reach at the flyover observer is 80 ft lower than the 737, resulting in a noise increase of perhaps 

0.4 EPNdB. The N+1 airplane’s lower acceleration reduces flyover airspeed (by 16 ktas) and contributes 

to an increased flyover EPNL duration component. 

If the trajectory and throttle settings had been assumed from previous experience rather than 

calculated, the effects noted above may result in under predicting each of the three EPNLs. If 737 

trajectory and throttle settings had been used for the N+1 UHB airplane, its certification noise would have 

been underestimated by nearly 2 cumulative EPNdB. 

The computed trajectories and throttle settings are used to determine engine and aircraft state 

information. This information in turn is used to predict component source noise spectra as described 

above for half-second time intervals along the trajectory. The spectra are propagated, and the OASPL and 

PNLT noise-time histories are plotted in Figures 14, 15, and 16 for the approach, lateral, and flyover 

observers, respectively. Observer time relative to the point of brake release (or touchdown) is used as the 

independent parameter in each figure. ANOPP has the ability to compute each noise source separately 

from the others. Plotting the levels of each source as a function of time provides additional insight to the 

overall problem. The OASPL metric is shown at the left in each figure because of its simplicity and its 

ability to clearly show the smooth rise and fall of each noise source over time. The PNLT metric—shown 

at the right in each figure—has qualities that capture level, frequency weighting, and tone annoyance 

penalties. Its time histories are therefore much more irregular than the OASPL histories: as the airplane 

approaches and recedes, Doppler and convective amplification effects have an effect on the PNLT 

metric’s frequency-weighting and tone penalties. The PNLT-time histories within the integration region 

of 10 PNdB from the maximum PNLT (shown as a horizontal line in each right-hand figure) are the 

regulatory bases for the EPNLs shown in Figure 17. 

The noise at the approach observer seems to be clearly dominated by the several airframe noise 

sources on an OASPL basis (Fig. 14, left), with core noise apparently making a significant contribution as 

well. However, when the higher-frequency fan noise and fan interaction tone contributions are considered 

via the PNLT metric (Fig. 14, right), fan noise becomes the dominant approach source. Another reason 

for the fan’s strong showing is that the acoustic signature of the Advanced Ducted Propulsor “Fan 1” is 

dominated by strong aft-quadrant noise levels (see Ref. 23.). This results in two PNLT peaks within the 

area of integration. The first peak consists of fore-quadrant fan noise increased by convective 

amplification, while the second, aft-quadrant fan noise peak rises despite the favorable convection effects 

of the receding aircraft. The strong aft fan noise extends the PNLT trace for several seconds beyond the 

airframe noise sources and increases the approach EPNL duration component. The airframe noise 

sources—effectively muted by their noise reduction technologies—make a secondary contribution to the 

approach EPNL. Core noise becomes a minor contributor on a PNLT basis, and jet noise at the approach 

power setting is insignificant relative to the other sources with any metric. The overall approach EPNL is 

86.2 EPNdB. 

The noise at the lateral observer is dominated by jet noise, and—perhaps surprisingly—core noise, at 

least on an OASPL basis (Fig. 15, left). But fan noise again rises to prominence on a PNLT basis when 

discrete interaction tones and spectral content are considered (Fig. 15, right). Fan noise is once again 

enhanced due to its strong aft-quadrant noise and twin PNLT peaks within the lateral observer area of 

integration. Jet noise and core noise, owing to their low-frequency, broadband content, are favored by the 

PNLT metric. Even so, the levels of lateral core noise are notable, rivaling even fan noise. Somewhat 
NASA/TM—2010-216085 14 

problematic—as discussed earlier—is the core noise prediction method’s inability to predict core noise at 

overall pressure ratios in excess of 30. The overall pressure ratio at lateral power and 180 ktas is 

approximately 34, resulting in perhaps an under prediction of lateral core noise by an unknown amount. 

The contribution of jet noise to the lateral EPNL is very small. If the core nozzle chevrons are analytically 

removed, lateral noise is predicted to increase by only 0.1 EPNdB. This suggests that—for this design—

the core nozzle chevrons may be omitted if their thrust performance penalty is unacceptable. The 

contributions of the airframe noise sources to the lateral EPNL are negligible. Gear are retracted after use, 

and the slat gaps are effectively sealed on their departure setting, resulting in no contribution to lateral 

noise at all. Flap and trailing edge noise is present, but makes no significant contribution to the lateral 

EPNL. The overall lateral EPNL is 81.5 EPNdB. 

Figure 14.—Approach observer OASPL noise-time histories (left), and PNLT noise-time histories (right). 

76

78

80

82

84

86

88

90

92

94

-45 -40 -35 -30 -25

Observer Time, s

Approach PNLT, PNdB

64

66

68

70

72

74

76

78

80

82

84

-45 -40 -35 -30 -25

Observer Time, s

Approach OASPL,dB

Fan

Jet

Core

All Airframe

Total

10 PNdB

Fan

Jet

Core

All Airframe

Total

10 PNdB

Fan

Jet

Core

All Airframe

Total

Figure 16.—Flyover observer OASPL noise-time histories (left), and PNLT noise-time histories (right). 

72

74

76

78

80

82

84

86

88

85 90 95 100 105 110

Observer Time, s

Flyover PNLT, PNdB

66

68

70

72

74

76

78

80

85 90 95 100 105 110

Observer Time, s

Flyover OASPL, d

B

Fan

Jet

Core

All Airframe

Total

10 PNdB

Fan

Jet

Core

All Airframe

Total

10 PNdB

Fan

Jet

Core

All Airframe

Total

Figure 15.—Lateral observer OASPL noise-time histories (left), and PNLT noise-time histories (right). 

70

72

74

76

78

80

82

84

86

60 65 70 75 80 85

Observer Time, s

Lateral PNLT, PNdB

66

68

70

72

74

76

78

60 65 70 75 80 85

Observer Time, s

Lateral OASPL, dB

Fan

Jet

Core

All Airframe

Total

10 PNdB

Fan

Jet

Core

All Airframe

Total

10 PNdB

Fan

Jet

Core

All Airframe

Total
NASA/TM—2010-216085 15 

The noise at the flyover observer has significant PNLT contributions of aft fan, core, and even 

airframe trailing edge noise (Fig. 16, right, and Fig. 17). The throttle cutback effectively eliminates jet 

noise altogether. The overall pressure ratio at the noise abatement cutback throttle setting is less than the 

core noise model maximum, so core flyover noise may reasonably be expected to be accurate, according 

to early results from the acoustic tool benchmarking study (Ref. 16). That core noise contributes at all is a 

source of concern, however, as is the unexpectedly high level of airframe trailing edge noise. The overall 

flyover EPNL is 81.6 EPNdB. 

The margins relative to the Stage 3 and 4 noise rule stringency limits are computed based on an 

airplane having a maximum takeoff gross weight of 159,200 lb (See Ref. 5.). The approach, lateral, and 

flyover EPNLs computed for the N+1 airplane are plotted in Figures 18, 19, and 20, respectively. Also 

plotted in the figures are the published EPNLs of all currently-certificated 737 models equipped with 

CFM56-7B series turbofans, as well as the maximum takeoff gross weight-dependent Stage 3 noise limits 

(defined in Ref. 1 and denoted by red lines). 

 Figure 19.—Lateral certification noise. 

80

82

84

86

88

90

92

94

96

98

100 1000 Maximum Takeoff Gross Weight, 1000 lb

Lateral EPNL, EPNdB

Stage 3 limit

All certificated 737s

equipped with CFM56-7Bs

N+1 airplane; 81.5 EPNdB,

-15.2 EPNdB below Stage 3

Figure 18.—Approach certification noise. 

84

86

88

90

92

94

96

98

100

100 1000

Maximum Takeoff Gross Weight, 1000 lb

Approach EPNL, EPNdB

All certificated 737s

equipped with CFM56-7Bs

Stage 3 limit

N+1 airplane; 86.2 EPNdB,

-14.2 EPNdB below Stage 3

Figure 17.—Predicted N+1 airplane certification noise levels. 

55

60

65

70

75

80

85

90

Fan

Jet

Core

Total Propulsion

Main Gear

Nose Gear

Slats

Flaps

Trailing Edge

Total Airframe

Total

EPNL, EPNdB

Approach

Lateral

Flyover
NASA/TM—2010-216085 16 

The FAA’s new Stage 4 noise regulations are based on the Chapter 4 noise standard approved by 

ICAO in 2001 in Annex 16 of their International Standards and Recommended Practices. The Stage 4 

noise standard applies to any new subsonic aircraft type application submitted after January 1, 2006, for 

countries that use Annex 16 as its noise certification basis. Loosely stated, the FAA interprets Stage 4 

limits to be: 1) the airplane may not exceed the Stage 3 limits at any of the three measurement points 

(Stage 3 rules allowed small trades); 2) the cumulative margin relative to the Stage 3 limits must be 

10 EPNdB or greater; and 3) the sum of any two of the Stage 3 margins must be 2 EPNdB or greater. 

The N+1 airplane’s computed margins relative to the Stage 3 limits for the approach, lateral, and 

flyover observers are -14.2, -15.2, and -9.7 EPNdB, respectively. The predicted cumulative margin 

relative to Stage 3 is –39 EPNdB; making the cumulative Stage 4 margin –29 EPNdB. The N+1 airplane 

has considerable margin at the approach and lateral locations, especially when compared to the 

737/CFM56-7B models currently in service. The margin at the flyover location, however, is less, and 

even lies within the lower range of the 737s. This is attributed to the behavior of the Advanced Ducted 

Propulsor Fan 1: its overall noise does not drop significantly when rotational speed is reduced from 

maximum to the cutback condition. In the future, improved design practices for UHB fans may more 

greatly reduce fan noise at cutback power and improve the flyover margin. 

Conclusions 

An in-depth, self-consistent, systems analysis study of an engine and airplane system using accepted 

NASA tools and methods is presented. Progress is measured toward the NASA Subsonic Fixed Wing 

Project’s N+1 noise goal of -32 cumulative EPNdB relative to Stage 4 limits. This analysis places the 

notional N+1 airplane’s Stage 4 cumulative margin at –29 EPNdB; or 3 EPNdB short of the earlier 2005 

study that established the goal. To add perspective, the Stage 4 cumulative margins of current 737-600, 

–700, –800, and –900 models range between –1 to –8 EPNdB, making the N+1 airplane considerably 

quieter than its predecessors in that vehicle class. As another point of reference, the new PurePower 

PW1000G 17,000 to 23,000 lb thrust class geared turbofan, slated for the CSeries twinjet, is currently 

projected by Pratt & Whitney to have a cumulative Stage 4 margin of –20 EPNdB (Ref. 36). 

The 3 EPNdB discrepancy between the –29 EPNdB margin calculated in this assessment and the 

earlier 2005 study that established the –32 EPNdB goal may be explained by any or all of the following: 

1) the incorporation of new thrust lapse, trajectory, and throttle modeling; 2) the unexpectedly high level 

of fan noise at the flyover condition; and 3) the use of improved analytical tools, methods, and 

assumptions. Despite this new result, the –32 EPNdB N+1 noise reduction goal remains the same. This 

3 EPNdB shortfall indicates that a slightly greater reliance on noise reduction technologies will be needed 

to achieve the goal. Some technologies under consideration that may be used to close the gap may include 

scarfed inlets, low-count stator technologies, and advanced acoustic liners, such as 3-degree of freedom, 

zero-splice, or nose lip liners. 

Figure 20.—Flyover certification noise.

79

81

83

85

87

89

91

93

100 1000 Maximum Takeoff Gross Weight, 1000 lb

Flyover EPNL, EPNdB

Stage 3 limit (twinjets)

All certificated 737s

equipped with CFM56-7Bs

N+1 airplane; 81.6 EPNdB,

-9.7 EPNdB below Stage 3
NASA/TM—2010-216085 17 

As shown by this study, as the once-historically dominant noise sources such as jet noise and fan 

noise are pushed lower by way of ultrahigh bypass ratio engine cycle design and innovative noise 

reduction technologies, other noise sources become important and require attention. For example, 

airframe trailing edge noise is seen here to perhaps now be a contributor to flyover noise. And engine core 

noise, in this case exacerbated by an engine design incorporating a high overall pressure ratio, contributes 

to lateral and flyover noise. A need is foreseen for improved engine core noise modeling methods at 

NASA that reflect the high overall pressure ratios of modern engines. 

Lastly, this airplane represents the “low-noise corner” of the design trade space, and it may not be 

characteristic of a balanced, profitable, business-case airplane design. In the interest of noise reduction, it 

has an engine cycle design that would not result in minimum block fuel, takeoff gross weight, or cost. 

This design also uses highly advanced engine and airframe noise reduction technologies to come within 

reach of NASA’s aggressive goal. An actual 737 replacement aircraft may not be as quiet as the N+1 

vehicle studied here. However, this study does indicate that if an aggressive, noise-focused design is 

pursued, it is possible to come within reach of the noise goal. 

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airworthiness certification. 

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Fixed Wing N+1 Noise Goal Study,” Acoustics Technical Working Group Meeting, Williamsburg, 

VA, Sep. 23–24, 2008. 

3. Envia, E., Berton, J.J., Bridges, J.E., Burley, C.B., Jones, M.G., Khorrami, M.R., and Rawls, J.R.: 

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16. Dahl, Milo D.: “A Process for Assessing NASA's Capability in Aircraft Noise Prediction 

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17. Burley, C.L.; Rawls, J.W., Jr.; Berton, J.J.; and Marcolini, M.A.: “Aircraft System Noise Prediction,” 

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Static%20Files/Docs/commercial_purepower_july08.pdf [cited 3/19/2009]
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4. TITLE AND SUBTITLE

An Analytical Assessment of NASA’s N+1 Subsonic Fixed Wing Project Noise Goal 

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Berton, Jeffrey, J.; Envia, Edmane; Burley, Casey, L. 

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WBS 561581.02.08.03.13.03 

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National Aeronautics and Space Administration 

John H. Glenn Research Center at Lewis Field 

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14. ABSTRACT

The Subsonic Fixed Wing Project of NASA’s Fundamental Aeronautics Program has adopted a noise reduction goal for new, subsonic,

single-aisle, civil aircraft expected to replace current 737 and A320 airplanes. These so-called “N+1” aircraft--designated in NASA 

vernacular as such since they will follow the current, in-service, “N” airplanes--are hoped to achieve certification noise goal levels of 32 

cumulative EPNdB under current Stage 4 noise regulations. A notional, N+1, single-aisle, twinjet transport with ultrahigh bypass ratio 

turbofan engines is analyzed in this study using NASA software and methods. Several advanced noise-reduction technologies are 

empirically applied to the propulsion system and airframe. Certification noise levels are predicted and compared with the NASA goal. 

15. SUBJECT TERMS

Jet aircraft noise; Noise reduction; Noise prediction (aircraft); Aeroacoustics 

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