Christopher E. Hughes, Gary G. Podboy, Richard P. Woodward, and Robert J. Jeracki

Glenn Research Center, Cleveland, Ohio

The Effect of Bypass Nozzle Exit Area on Fan 

Aerodynamic Performance and Noise 

in a Model Turbofan Simulator

NASA/TM—2013-214029

December 2013

GT2005–68573
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Christopher E. Hughes, Gary G. Podboy, Richard P. Woodward, and Robert J. Jeracki

Glenn Research Center, Cleveland, Ohio

The Effect of Bypass Nozzle Exit Area on Fan 

Aerodynamic Performance and Noise 

in a Model Turbofan Simulator

NASA/TM—2013-214029

December 2013

GT2005–68573

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

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Level of Review: This material has been technically reviewed by technical management. 
NASA/TM—2013-214029 1

The Effect of Bypass Nozzle Exit Area on Fan Aerodynamic 

Performance and Noise in a Model Turbofan Simulator

Christopher E. Hughes, Gary G. Podboy, Richard P. Woodward, and Robert J. Jeracki

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Abstract

The design of effective new technologies to reduce aircraft 

propulsion noise is dependent on identifying and 

understanding the noise sources and noise generation 

mechanisms in the modern turbofan engine, as well as 

determining their contribution to the overall aircraft noise 

signature. Therefore, a comprehensive aeroacoustic wind 

tunnel test program was conducted called the Fan Broadband 

Source Diagnostic Test as part of the NASA Quiet Aircraft 

Technology program. The test was performed in the anechoic 

NASA Glenn 9- by 15-Foot Low Speed Wind Tunnel using a 

1/5 scale model turbofan simulator which represented a 

current generation, medium pressure ratio, high bypass 

turbofan aircraft engine. The investigation focused on 

simulating in model scale only the bypass section of the 

turbofan engine. The test objectives were to: identify the 

noise sources within the model and determine their noise 

level; investigate several component design technologies by 

determining their impact on the aerodynamic and acoustic 

performance of the fan stage; and conduct detailed flow 

diagnostics within the fan flow field to characterize the 

physics of the noise generation mechanisms in a turbofan 

model. This report discusses results obtained for one aspect 

of the Source Diagnostic Test that investigated the effect of 

the bypass or fan nozzle exit area on the bypass stage 

aerodynamic performance, specifically the fan and outlet 

guide vanes or stators, as well as the farfield acoustic noise

level. The aerodynamic performance, farfield acoustics, and 

Laser Doppler Velocimeter flow diagnostic results are 

presented for the fan and four different fixed-area bypass 

nozzle configurations. The nozzles simulated fixed engine 

operating lines and encompassed the fan stage operating 

envelope from near stall to cruise. One nozzle was selected 

as a baseline reference, representing the nozzle area which 

would achieve the design point operating conditions and fan 

stage performance. The total area change from the smallest to 

the largest nozzle was 12.9 percent of the baseline nozzle 

area.

The results will show that there are significant changes in 

aerodynamic performance and farfield acoustics as the fan 

nozzle area is increased. The weight flow through the fan 

model increased between 7 and 9 percent, the fan and stage 

pressure dropped between 8 and 10 percent, and the adiabatic 

efficiency increased between 2 and 3 percent—the magnitude 

of the change dependent on the fan speed. Results from force 

balance measurements of fan and outlet guide vane thrust 

will show that as the nozzle exit area is increased the 

combined thrust of the fan and outlet guide vanes together 

also increases, between 2 and 3.5 percent, mainly due to the 

increase in lift from the outlet guide vanes. In terms of 

farfield acoustics, the overall sound power level produced by 

the fan stage dropped nearly linearly between 1 dB at takeoff 

condition and 3.5 dB at approach condition, mainly due to a 

decrease in the broadband noise levels. Finally, fan swirl 

angle survey and Laser Doppler Velocimeter mean velocity 

and turbulence data obtained in the fan wake will show that 

the swirl angles and turbulence levels within the wake 

decrease as the fan nozzle area increases, which helps to 

explain the drop in the fan broadband noise at all fan speeds.

Nomenclature

A Cross sectional area, in2

c Airfoil chord, in

M Mach number

N Mechanical fan rotational speed, rpm

Nc Corrected fan speed, 

θ

N

NPc Percent of corrected fan design speed, 100

N

N

dp

c ⋅ ,

percent

P Pressure, psia

R Gas constant, 53.35 ft lbf/lbm°R

T Temperature, °R

t Airfoil thickness, in

γ Specific heat ratio, 1.4

δ Pressure correction to standard day conditions,

14.696

Pt,o

η Adiabatic efficiency, 

1 Tt,o

Tt

1

1

Pt,o

Pt

 − 

 





− γ

γ−





 





θ Temperature correction to standard day conditions, 

518.67

Tt,o
NASA/TM—2013-214029 2

ω Weight flow rate,

2( 1)

1

2

bm

bm

t,bm

t,bm

bm

M

2

1 1

M

T

P

R

A

γ−

γ+





 



 γ − +

γ , lbm/sec

ωc Corrected weight flow rate, δ

θ

ω , lbm/sec

Θ Acoustic angle, deg

Subscripts

ae Adiabatic efficiency

bm Bellmouth inlet condition

c Corrected condition

dp Design point

em Emission

f Fan value

f Force

geom Geometric

m Mass

o Freestream condition

s Static condition

s Stage value

t Total condition

Introduction

In recent years, commercial aircraft noise has become a 

major concern for aircraft owners and airport operators. The 

increased frequency of takeoffs and landings has produced an 

increasing number of complaints from local residents. The 

Federal Aviation Administration in the United States and the 

International Civil Aviation Organization, the international 

organization that coordinates environmental noise issues, 

have responded to these complaints by issuing increasingly 

more stringent noise regulations and curtailed flight 

operations for aircraft, forcing aircraft and engine 

manufacturers to pursue quieter aircraft designs. With the 

support of Congress, NASA and the U.S. aircraft and engine 

manufacturing companies have joined to cooperatively 

investigate high-risk technologies for reducing aircraft noise 

through programs such as the NASA Advanced Subsonic 

Technology (refs. 1 to 4) and Quiet Aircraft Technology 

Programs.

As part of an overall NASA effort to reduce aircraft noise, 

technical programs were initiated starting in 1989 with the 

major U.S. aircraft manufacturers to investigate noise 

reduction technologies. As part of this effort, NASA 

established aggressive goals to reduce the noise signature of 

then 1992 technology turbofan engines by 6 EPNdB

(Effective Perceived Noise dB) by the year 2000. Engine 

studies were conducted across a wide range of engine and 

aircraft operating cycles to identify and quantify the benefit 

of potential noise reduction concepts. Several noise reduction 

technology concepts were investigated using scale model 

wind tunnel testing of turbofan engine simulators, and the 

noise reduction potential successfully demonstrated in most 

cases (refs. 5 to 10). However, new noise reduction standards 

and new noise reduction program goals were aggressively 

pushing the technology. The goals of the NASA Quiet 

Aircraft Technology Program, the follow-on the Advanced 

Subsonic Technology program, sought to reduce turbofan 

noise by another 4 EPNdB by 2006, using 1997 High Bypass 

turbofan engine technology as the baseline. This new level of 

noise reduction technology would require novel technical 

approaches.

Therefore, in order to more fully understand the noise 

sources and noise generation mechanisms in a modern 

turbofan engine and be able to properly guide further noise 

reduction technology development, a comprehensive scale 

model wind tunnel test of a turbofan simulator was planned,

called the “Fan Broadband Source Diagnostic Test,” or just 

Source Diagnostic Test. The test was a cooperative effort 

between NASA and General Electric Aircraft Engines. It was 

a two-phase experimental investigation conducted in the 

NASA Glenn anechoic 9- by 15-Foot Low Speed Wind 

Tunnel to identify and understand the noise source 

mechanisms within a turbofan engine and determine their 

individual contributions to the overall engine system noise 

signature. For this test, the bypass stage portion of a medium 

pressure ratio, high bypass ratio turbofan engine was 

simulated in approximately 1/5 model scale. The fan model 

consisted of a 22 in., 22-blade, wide-chord fan, an outlet 

guide vane or stator assembly, and a simulated flight-type 

nacelle which included a fixed area nozzle and an inner 

flowpath contour simulating the outer core cowling. The 

emphasis for this test was placed only on simulating the 

bypass stage portion of the engine, not the booster core or 

power stage, eliminating the possibility of contaminating the 

fan noise field with the noise from a simulated core. 

Likewise, to minimize the noise contamination sources 

within the fan model, the outlet guide vanes were used to 

support the nacelle, eliminating any struts or pylons from the 

flowpath. The fan model was powered by the NASA Glenn 

Ultra High Bypass Drive Rig propulsion simulator. In order 

to simulate aircraft flight effects during takeoff roll, approach 

and landing phases, wind tunnel velocities up to Mach 

number 0.10 were provided during acoustic testing.

The Source Diagnostic Test had several technical 

objectives, both acoustic and aerodynamic. The first phase of 

testing, completed in 2000, verified the usability of the 

proposed fan as a baseline for future technology 

development. The aerodynamic performance and farfield 

acoustics of the fan were measured. In addition, the effect of 

the number of outlet guide vanes used, as well as the effect of 

aft radial sweep and chord length, on the aerodynamic 

performance and noise level of the fan model was 

investigated with three different outlet guide vane designs 

(ref. 11). The overall test had several areas of investigation 

including acoustic mode measurements using sensors located 

on the inner surface of the fan duct, spinning mode 
NASA/TM—2013-214029 3

measurements using a rotating acoustic rake in both the inlet 

and the nozzle, unsteady surface pressure measurements on 

two different types of outlet guide vanes, and detailed flow 

diagnostic measurements using laser doppler velocimetry and 

hot-wire anemometry. Some results from this phase of the 

test can be found in references 12 to 16. Also, the first rotor 

alone testing in a wind tunnel using realistic fan geometry 

was accomplished (ref. 14) This part of the testing was 

important for understanding the contribution of the fan noise 

sources to the overall engine system noise.

In phase two, completed in 2003, the same fan model 

components (including the baseline fan and outlet guide 

vanes (OGVs)) were used to determine the effect of fan tip 

clearance and bypass nozzle exit area on the aerodynamic 

performance and noise. Previous research has been 

conducted to determine the effect of nozzle area on fan 

performance and noise (refs. 17 to 21). However, this work 

was performed using then current technology that, while not 

directly applicable to current generation turbofans, did 

provide invaluable direction and guidance for future areas of 

noise reduction research. In addition, the previous 

investigations were conducted using a more component riglike research apparatus. The results presented in this paper 

have a direct relationship with current technology, which can 

be scaled directly to full scale application. In addition, a 

comprehensive set of flow diagnostic surveys using Laser 

Doppler and Particle Image Velocimetry were performed to 

measure the unsteady velocity and turbulence components 

within the internal model flowfield. These data allow a better 

understanding of the flow physics and how it interacts with 

the model hardware, leading to identifying and understanding 

the mechanisms that produce noise within the model.

This paper will discuss the results of an investigation to 

determine the effect of the bypass nozzle exit area on the fan 

bypass stage aerodynamic and acoustic performance. Four 

different nozzle area geometries were designed and 

fabricated for use with a baseline fan and outlet guide vanes.

The total area change investigated was 12.9 percent of the 

baseline nozzle area, which was the original Source 

Diagnostic Test Phase One fan nozzle, designed to achieve 

optimum fan stage performance at the design point 

conditions. Relative to the baseline nozzle, one smaller 

nozzle and two larger nozzle areas were tested. The 

aerodynamic performance of the fan and OGVs with each of 

the four nozzle configurations will be discussed in terms of 

fan and stage pressure, temperature and adiabatic efficiency 

maps including radial profiles downstream of the fan.

Farfield acoustics for the fan model with three nozzle 

configurations (the baseline nozzle and the two larger area 

nozzles) will be provided in terms of reduction of overall 

sound power level noise at various fan speeds. Finally, flow 

diagnostics will be shown which will describe the mean 

velocity and turbulence components in the fan wake for three 

of the nozzles for several fan speeds in order to show the 

changes in these components with increasing nozzle area and 

their relationship to the farfield acoustics.

Research Apparatus

NASA Glenn Research Center 9- By 15-Foot Low 

Speed Wind Tunnel

The 9- by 15-Foot Low Speed Wind Tunnel (9x15) is an 

anechoic wind tunnel facility located at the NASA Glenn 

Research Center in Cleveland, Ohio. The facility is operated 

as an open loop, continuous flow wind tunnel at atmospheric 

pressure conditions. The test section is located in the return 

leg of the NASA Glenn 8- by 6-Foot Supersonic Wind 

Tunnel flow circuit. The two wind tunnels share a common 

airflow drive system. The 9x15 wind tunnel is capable of 

producing Mach numbers 0.0 to 0.23 (ref. 22) in the test 

section. Static propulsion system performance testing is also 

conducted in this facility. Flow conditioning upstream of the 

test section allows the facility to produce very low freestream 

turbulence and distortion levels, making it ideal for acoustic 

testing of propulsion systems (ref. 23). The test section 

surfaces are covered with boxes filled with an acoustic 

treatment material that is capable of absorbing sound 

reflections down to 250 Hz (refs. 24 to 26). Figure 1 is an 

overview of the wind tunnel circuit. Figures 2a and b provide 

an overhead view drawing of the 9x15 test section and the 

location of the installed research hardware and 

instrumentation for aerodynamic performance and farfield 

acoustic configurations. 

Turbofan Propulsion Simulator

A propulsion simulator called the NASA Glenn Ultra High 

Bypass (UHB) Drive Rig was used to power the model fan 

test article. Details about the UHB Drive Rig can be obtained 

from a report documenting the General Electric Aircraft 

Engines (GEAE) Universal Propulsion Simulator (ref. 27), 

which is very similar to the NASA Glenn UHB Drive Rig 

simulator. A four-stage air turbine generates the power that is 

supplied to the fan model through a common shaft 

connection. The air turbine is driven by high pressure (up to 

350 psi), high temperature (up to 550 °F) air that is supplied 

to it from tubes running through a support strut that mounts 

the UHB Drive Rig in the wind tunnel test section. The UHB 

Drive Rig can generate a maximum of 5,000 shaft 

horsepower at 16,850 rpm. Figure 3 provides a cutaway view 

of the UHB Drive Rig and a table of maximum performance 

parameters.

Fan Module

The research model, or fan module, used was a 1/5-scale 

model representation of the bypass stage of a current 

generation high bypass turbofan aircraft engine. The fan 

module was designed and built by GEAE with partial 

funding under contract to NASA Glenn. Only the bypass 

section of the engine was simulated in order to ensure that 

the fan model noise field was not contaminated by noise from 

a core section simulator. The fan module consisted of the fan, 
NASA/TM—2013-214029 4

the outlet guide vanes (OGVs) and a flight-type nacelle. The 

nacelle included a flight-type inlet, a cowl and a fixed-area, 

flight-type bypass exhaust nozzle. In order to minimize 

adding additional noise sources within the model, the OGVs 

were designed to provide structural support for the nacelle, 

thereby eliminating the need for struts normally present in a 

turbofan engine. In addition, neither pylon nor bifurcations 

were simulated in the model. Figure 4 shows a cutaway view 

of the fan module in the flight configuration, showing the 

location of various model components.

The fan used for this test was 22 in. in diameter and had 

22 individual, wide-chord, titanium blades. In combination 

with the 54 vane baseline OGVs, the fan model had a stage 

design point pressure ratio of 1.47 at a model corrected speed 

of 12,657 NPc, corresponding to a design point fan tip speed 

of 1,215 ft/s. Table 1 provides a summary of the design 

parameters for the fan. The fan was a scale model designed 

and previously tested by GEAE, who designated the fan as 

“R4.” The fan was originally designed to operate in

conjunction with a powered core simulator. As a result, the 

performance level at its design point could not be achieved in 

this test since this installation did not include a core 

simulation. However, since this fan was meant to be 

representative of current technology, the performance 

compromise was deemed to be acceptable for this test. The 

fan was tested with a rubstrip casing that was designed for a 

.020 in. fan blade tip clearance at the fan design point (100% 

corrected fan speed, or 12,657 NPc). This clearance was 

selected as representative of a turbofan engine which had 

undergone many takeoff and landing cycles, and therefore 

could potentially be the noisiest configuration. In addition, 

this tip clearance minimized the chance of a fan rub event 

during testing, thus insuring a clean and uniform flowpath 

contour at the fan tip. This was important for obtaining 

uncontaminated acoustic data and LDV fan wake flow 

diagnostic survey data, since a discontinuity in the flowpath 

after a rub would alter the tip flowfield.

There were four fixed-area fan nozzles designed for 

testing. Typically, fixed-area, “flight-type” engine nozzles 

are sized so the engine can achieve maximum efficiency at 

cruise conditions. For this test, a baseline nozzle and three 

additional fan nozzles were included which allowed the fixed 

operating line to be varied across the fan and stage operating 

maps. The purpose of the additional fan nozzles was to 

determine the effect of reducing the fan blade loading on the 

fan module acoustics. The four fan nozzles were: 1), a 

Baseline (BL) nozzle, typically used as the nozzle for all 

acoustic testing, and designed to achieve minimum pressure 

losses across the OGVs and maximum efficiency at the 

design point (normally, cruise conditions); 2), a Low Flow

(LF) nozzle, which reduced the fan exit area by 2 percent

compared with the BL nozzle and the maximum fan weight 

flow by 2.5 percent at the fan design speed; 3), a Design 

Point (DP) nozzle, which increased the fan exit area 

5.4 percent and the maximum weight flow by 5 percent; and 

4), a High Flow (HF) nozzle that increased the fan exit area 

10.9 percent and the maximum weight flow by 7.5 percent.

The size for the additional nozzles was determined using 

previous fan performance results (ref. 11) and isentropic flow 

relationships in an iterative spreadsheet program. A plot of 

the four flight-type fan nozzle geometries is shown in 

figure 5. Table 3 lists the design parameters for the four fan 

nozzles.

To establish the fan and OGV performance, the fan 

module installation included a uniform-inflow bellmouth 

inlet and either the fixed-area flight-type nozzle or a Variable 

area Fan Exit Nozzle (VFEN). The fixed area nozzle was 

used to obtain the fan performance on a representative 

operating line for a turbofan engine installation, at sea level 

conditions. The VFEN was used to obtain fan and stage 

performance across a range of fan speed operating conditions 

and simulated aircraft flight conditions. It consisted of a 

series of trapezoidal-shaped plates, each with a central radial 

pivot, arranged circumferentially in an annular duct. The 

plates moved in pairs in opposing directions to one another, 

like double doors. The fan operating point was changed by 

varying the exit area, and therefore the back pressure on the 

fan and weight flow through the fan stage, while at a constant 

fan speed. Changing the fan back pressure simulates a change 

in the aircraft flight speed and altitude. In figure 6a, 

photographs of the fan module in the aerodynamic 

performance configuration while installed in the 9x15 wind 

tunnel test section are shown. Figure 6b is a close-up view of 

the VFEN arrangement.

Laser Doppler Velocimetry

In order to conduct the LDV wake survey it was necessary 

to place part of the LDV system inside the test section of the 

wind tunnel. Figure 7a shows a photograph of the LDV 

traverse system located on the side of the fan model. The 

traverse was used to move the LDV probe volume radially 

and axially relative to the model. The LDV system optics are 

located behind the cylindrical shield shown in the photo. This 

shield was installed to keep the tunnel flow from striking the 

optics.

Figure 7b shows a photograph taken with the cylindrical 

shield removed. In this photo the fiber optic cables used to 

deliver the laser beams into the tunnel, the transmitting optics 

used to direct the beams into the model, and one set of 

receiving optics can be seen. The LDV system is a fourbeam, two-color, backscatter system which allows the 

measurement of two components of velocity simultaneously.

Two green beams were used to measure the axial component 

of velocity, while two blue beams allowed the measurement 

of the tangential component. The photo provided in figure 7b

shows one of two optical arrangements used during the test.

As pictured, only one component of velocity could be 

measured at a time. Initially, another optical arrangement, 

one employing two sets of receiving optics (one above and 

one below the transmitting optics) was used to conduct the 

wake surveys. During the initial surveys it was possible to 
NASA/TM—2013-214029 5

measure both the axial and tangential velocity components 

simultaneously.

The LDV flow diagnostic wake surveys were conducted 

with bellmouth inlet installed on the model. The bellmouth 

inlet provided essentially the same flow into the fan as the 

flight inlet, but allowed the tunnel to be run at a lower speed.

This provided the following benefits: 1) the test was 

conducted at a lower cost and 2) the LDV system 

components mounted in the test section were subjected to 

lighter air loads. The test section Mach number during the 

LDV testing was approximately 0.05. Two windows installed 

in the side of the model permitted optical access to the 

internal flow. These two windows are shown in the 

photograph of figure 8b. The downstream window shown at 

the left was used to acquire the wake surveys. These 

windows, made of 0.1 in. thick sodium alumino silicate, were 

slumped in a furnace to have the same shape as the inner 

contour of the model.

In order to acquire LDV data at the axial location 

corresponding to the leading edge of the Baseline stators, a 

different set of OGVs were installed in the fan module. This 

alternate set consisted of 26 vanes, each with 30° of leading 

edge radial sweep. The increase in the axial length between 

the fan trailing edge and the OGV leading edge provided by 

the swept OGVs allowed a large downstream LDV window 

to be used, which in turn allowed the Baseline OGV leading 

edge axial location to be viewed by the LDV system. Test 

results previously obtained for the fan module during the

Source Diagnostic Test Phase 1 have shown that the fan and 

stage performance were relatively unaffected by replacing 

the Baseline OGVs with the swept OGVs (ref. 11).

Research Instrumentation

Aerodynamic 

Freestream conditions in the wind tunnel were determined 

using a ceiling mounted pitot-static rake with thermocouples 

located near the entrance to the test section. Fan inlet 

conditions to the bellmouth were determined using a floor 

mounted, cruciform-shaped rake located near the fan 

centerline and upstream of the bellmouth inlet. Total pressure 

and total temperature conditions upstream of the fan and 

within the fan stream tube were measured using this rake.

The fan weight flow was determined using static pressure 

measurements obtained within the bellmouth inlet and a flow 

correlation function which related the average of the 

bellmouth static pressures to the fan weight flow.

Fan and stage performance were determined using fixed 

total pressure/total temperature rakes mounted behind the fan 

and OGVs. Fan performance was obtained using three rakes 

and stage performance was obtained with seven rakes. Each 

rake consisted of seven measurement sensors, and each 

sensor contained a total pressure probe and a total

temperature probe co-located within an aspirated stagnation 

tube. The sensors on each rake were located radially in such a 

way as to provide flow conditions at the center of equal 

areas. In addition, surface mounted static pressures were 

located at several axial locations in the fan module for 

calculating internal velocities. During the fixed operating line 

testing with the fixed area nozzle installed, only fan 

performance could be measured, since the stage performance 

rakes could only be installed in the model with the VFEN in 

place. A complete description of the aerodynamic 

performance instrumentation used in this test can be found in 

reference 11.

Forces produced by the fan and OGV were measured 

using two different types of force balances. Thrust and torque 

produced by the fan was measured with a two-component 

rotating force balance. Signals from this balance were 

transmitted to the data system using a 104 channel slip ring 

system internal to the UHB Drive Rig. Thrust and drag forces 

produced by the OGVs (including the nacelle and inner 

flowpath hardware downstream of the fan) were measured 

using a six-component static force balance mounted directly 

on the UHB Drive Rig. Signals were sent directly to the data 

system through wiring channels located in a fairing mounted 

in front of the Drive Rig strut. Figure 4 shows the location of 

the balances and their physical relationship to the fan module 

components. The accuracy of the measured thrust is ±10 lbf, 

or ±0.25 percent of the full scale measurement range 

(4000 lbf) of the combined balances. A complete description 

of the force balance components and technique used in 

measuring performance of turbofan simulators in wind tunnel 

testing at Glenn using can be found in reference 28.

Acoustic 

Figure 8a is a photograph of the fan module in the 

acoustic, flight-type configuration installed in the 9x15. In 

figure 8b, the downstream fixed microphones and sideline 

traversing microphone probe can be seen to the left of the fan 

model.

Sideline acoustic data were acquired with a computercontrolled translating microphone probe and with three aft 

microphone assemblies mounted to the tunnel floor. The 

translating microphone probe acquired data at 48 sideline 

geometric angles from 27.2° to 134.6° relative to the fan 

rotor plane. The translating probe traverse was 89 in. from 

the fan rotational axis (about four fan diameters). A wallmounted microphone probe was placed at a reference 

location adjacent to the translating probe home position 

(134.6°, maximum aft travel). The three fixed microphone 

assemblies were mounted at the home axial position to 

acquire aft acoustic data at geometric angles of 140°, 150°, 

and 160°. The acoustic data were acquired through a digital 

computer system and stored for post-run analysis.
NASA/TM—2013-214029 6

Experimental Procedure

Aerodynamic Performance

Fan and stage aerodynamic performance maps were

obtained for the baseline R4 fan and with the 54 vane 

Baseline OGVs installed. In addition, performance on fixed 

operating lines was obtained with each of the four fixed area 

fan exit nozzle configurations. A Mach number of 0.05 was 

set in the test section during testing in order to provide 

uniform temperature and pressure distributions into the fan, 

and also to prevent the fan from ingesting vortices from the 

test section surfaces. To eliminate the day-to-day variations 

in pressure and temperature that affect the performance 

calculations, the fan and stage performance parameters were 

corrected to standard day pressure and temperature 

conditions, where required. To insure that data was acquired 

at steady state conditions, a 30 second settling time interval 

was maintained after each new fan operating condition was 

reached. In addition, pressure and temperature information 

from the data system was time averaged over a 10 second 

sampling.

Fan and Stage Mapping

Fan and stage performance mapping was conducted with 

the bellmouth inlet and the VFEN installed on the fan 

module. A fan speed range from 50 to 100 percent of the fan 

design speed was investigated. At each speed condition, the 

maximum weight flow was achieved at the maximum nozzle 

area with the VFEN fully open. To determine stage adiabatic 

efficiency, it was assumed that there was no loss in total 

temperature across the OGVs; this permitted the total 

temperature data from the fan performance rakes to be used 

in the calculations. This method is a more accurate 

determination of stage adiabatic efficiency since variations in 

temperature measurements between the fan rakes and the 

stage rakes are eliminated, thereby also eliminating the 

accompanying errors in the stage adiabatic calculation.

Overall values for the fan and stage performance were 

obtained by averaging the seven radial profile values for each 

performance parameter. 

In order to minimize the risk of damaging the fan blades 

caused by an unintentional hard fan rub, the fan stall region 

located at the lower fan weight flow conditions was 

intentionally avoided. With this fan design, an approaching 

stall condition was indicated by an increase in the fan blade 

stress. Therefore, the minimum fan weight flow was 

established when the fan blade stress measured with blade 

mounted strain gauges reached a predetermined limit, 

typically 25 percent of the fan material yield strength. This 

limit, however, varied with fan speed. The 87.5 percent speed 

line was unusually sensitive (higher blade stress at higher 

weight flow conditions), so the minimum weight flow 

boundary at this speed was larger than the minimum weight 

flow at the other fan speed lines. The shape of the maps had a 

distinct bend, or knee, because of this sensitivity.

Fixed Nozzle Operating Line

Fan performance on the fixed area operating lines at sea 

level conditions was obtained using the bellmouth inlet and 

the fixed area, flight-type, nozzles. However, only the fan 

performance could be obtained with the fixed area nozzle 

installed since the stage performance rakes could only be 

installed with the VFEN installed. Therefore, once the fan 

weight flow and fan operating parameters were established 

for the fixed nozzle operating line, the corresponding stage 

performance was obtained with the VFEN installed by 

adjusting the nozzle exit area to match the fixed nozzle fan 

pressure and weight flow at each corresponding fan speed. 

Farfield Acoustics

The farfield acoustic testing was conducted with only 

three different fixed area nozzles—BL, DP and HF. During 

aerodynamic performance testing, the data showed that the 

LF performance was not too different than the BL 

performance on the operating line. Therefore, the LF nozzle 

was not acoustically tested. All of the fan stage acoustic data 

were acquired at tunnel Mach number of 0.10, which was 

sufficient to achieve acoustic flight effect (ref. 29). Sideline 

data are presented in terms of emission angles. The emission 

angles are related to the geometric or observed angles by the 

relationship:

Θem = Θgeom – sin-1 (Mo sin Θgeom),

where Θem and Θgeom are, respectively, the emission and 

observed sideline angles, and Mo is the test section Mach 

number. At Mach 0.10 then, the observed angles for the 

sideline translating microphone probe range from 25° to 

130°, and the three fixed microphones measure aft observed 

angles of 136°, 147°, and 158°. This angular range was 

sufficient to define the sideline noise profile for this aftdominated fan stage for subsequent EPNL calculations.

Digital acoustic data were processed as constant 

bandwidth spectra. Spectra were acquired and averaged at 

each translating probe or fixed microphone position with 5.9 

and 59 Hz bandwidths. These constant bandwidth spectra 

were electronically merged and used to generate 1/3-octave 

spectra, using the 5.9 Hz bandwidth results for lower 1/3rd

octave frequencies, and the 59 Hz bandwidth results for the 

higher frequencies. Sound power level (PWL) spectra were 

calculated from the SPL spectra assuming spherical 

symmetry through the range of sideline data acquisition.

Possible noise contributions outside the sideline range were 

ignored.

Laser Doppler Velocimetry Surveys

Figure 9 shows the location at which LDV wake 

measurements were made relative to the fan module

hardware. The axial location was 3.12 in. downstream of the 

non-rotating position of the tip trailing edge. Wake data were 
NASA/TM—2013-214029 7

acquired at four corrected rotor speeds, 6,329, 7,808, 11,074, 

and 12,657 (corresponding to 50, 61.7, 87.5, and 100% of the 

fan design speed). The 7,808, 11,074, and 12,657 speeds 

correspond to nominal approach, cut-back, and take-off 

conditions for the R4 fan. The wake surveys were conducted 

to determine how the wake flows vary with rotor speed and 

axial location.

The tunnel flow was seeded with polystyrene latex (PSL) 

spheres that were manufactured at the NASA Glenn Research 

Center. The nominal size of the PSL spheres is estimated to 

be approximately 0.7 μm in diameter. The polystyrene 

spheres are diluted with ethanol and sprayed into the wind 

tunnel using a set of nine spray nozzles located 

approximately 80 ft upstream of the test section. The liquid 

solvent evaporates by the time it reaches the test section, 

leaving behind the solid spheres on which the LDV data is 

obtained.

The individual velocity measurements were sorted into 

circumferential bins around the rotor using shaft angle 

encoders fed with the once-per-revolution signal of the rotor.

These encoders segmented the 360° of rotor revolution 

occurring between two consecutive once-per-revolution 

pulses into 1100 bins of equal width (50 bins per blade 

passage). Each time a velocity measurement was made, the 

encoder output was sampled to determine the number of bins 

generated since the occurrence of the previous once-per-rev 

pulse. The velocity and corresponding bin number were then 

stored in the computer as a data pair.

Data were acquired at the survey measurement location 

over many rotor revolutions until either a preset number of 

measurements had been acquired on one of the two LDV 

channels, or until the maximum time allotted for the data 

acquisition had elapsed. On-line data plots were used to 

determine the number of measurements required to 

accurately resolve the flows occurring within the individual 

blade passages. In general, the higher the unsteadiness in the 

flow, the greater the number of measurements required to 

resolve the flow. On average, more than 40,000 velocity 

measurements per component were obtained at each

combination of measurement location and operating 

condition. A more complete description of the LDV system 

and technique as used for flow diagnostics in turbofan 

simulators during wind tunnel tests at Glenn can be found in 

reference 12.

Results and Discussion

Aerodynamic Performance

For the results presented in this section, the accuracy of 

the performance calculations is based on empirical 

observation and repeat data points. The accuracy of the data 

acquisition systems used during testing were ±0.002 psia for 

pressure and ±0.25 °F for temperature. However, the data 

systems were configured to provide time-averaged 

measurements at a high sample rate. For temperature and 

pressure, the data values are based on an average of ten, onesecond averages, with each one-second average based on the 

average of 20,000 samples. Therefore, the accuracy of the 

discrete performance points is higher than the results based 

on discrete data samples. For pressure ratio, the accuracy is 

±0.0003; for temperature ratio, the accuracy of the results is 

±0.001; and for adiabatic efficiency, the accuracy of the 

results is ±0.3 percent.

Fan performance maps are presented in figure 10. The 

corrected weight flow, fan total pressure ratio, total 

temperature ratio and adiabatic efficiency are shown in 

Figures 10a through d, respectively. Shown on each of the 

performance map plots is the operating line performance 

measured for the four fixed area nozzles—Low Flow, 

Baseline, Design Point and High Flow. As stated in an earlier 

section, the last data point on the far left on the plot does not 

represent the fan stall line. Instead, this minimum weight 

flow condition data was obtained at what was considered a 

safe operating distance away from the fan stall line in order 

to avoid a fan tip rub or any possible damage to the fan 

blades due to high blade stress caused by unintentionally 

entering a fan stall condition. The results show that opening 

the nozzle area allows more weight flow through the model 

on each fan speed line (fig. 10a), moving the operating lines 

toward where the engine cruise operating line would be at the 

right side of the plots in the figure. Consequently, the 

pressure and temperature ratios decrease along each fixed fan 

speed line (figs. 10b and c). The reverse is true as the nozzle 

exit area is closed. The weight flow decreases, the fixed 

operating line moves toward the left side or stall, and the fan 

pressure and temperature ratios increase on each fan speed 

line. Using the BL nozzle data as a reference, the changes in 

pressure ratio from LF to HF nozzle ranged from 1.1 to 

–7.3 percent at 50 percent speed and 1.1 to –11.4 percent at 

100 percent speed. For the same LF to HF nozzle area range, 

the temperature ratio changed from 1.6 to –9.1 percent at 

50 percent speed and 1.4 to –9.4 percent at 100 percent

speed. The weight flow, meanwhile, changed from –1.4 to 

7.9 percent at 50 percent speed and –1.1 to 5.7 percent at 

100 percent speed. As can be seen, the pressure and 

temperature changes with nozzle area get larger as the fan 

speed increases, reflecting the spreading out of the fixed 

nozzle operating lines.

Interestingly, the fan adiabatic efficiency goes up as the 

nozzle area increases (fig. 10d), rising between 2.9 percent

(50% fan speed) and 1.9 percent (100% fan speed). Peak 

efficiency ranged from 91.1 percent for the LF nozzle to 

93.0 percent for the DP nozzle. The HF nozzle reached the 

same peak efficiency as the DP nozzle, but at 87.5 percent

fan speed. At 100 percent corrected fan speed, the HF nozzle 

pressure ratio, temperature ratio and adiabatic efficiency 

rapidly fall off. The adiabatic efficiency quickly drops about 

2 to 91.0 percent compared to the DP nozzle performance at 

the same speed. The HF nozzle area is too far open at this 

point, causing the performance fall-off. Since the fan tip is 

supersonic at this operating point (the fan tip reaches 
NASA/TM—2013-214029 8

transonic at the 87.5 percent corrected fan speed, or cutback, 

condition), the drop in performance is most likely due to the 

low blade incidence angle experienced at this high axial 

velocity condition, a result of the high weight flow into the 

fan. Another possibility is separated flow on the fan blades, 

caused by the interaction of the strong passage shock with the 

boundary layer on the blade. The passage shock at this 

condition would be stronger compared with the shocks at the 

same speed for the other nozzle areas because of the higher 

relative velocity at the fan tip as a result of the higher weight 

flow into the fan. Also, the corrected weight flow appears to 

level out somewhat at 100 percent fan speed for the HF 

nozzle case (fig. 10a), suggesting that the presence of 

stronger shocks at the outer portions is limiting the amount of 

flow passing through the fan.

The stage performance plots are shown in figure 11. As 

discussed earlier, the assumption is made that there are 

minimal total temperature losses across the OGVs, so the fan 

exit total temperatures are used in the calculation of the stage 

adiabatic efficiency. The same performance trends observed 

for the fan can be seen in these results, indicating that there 

are no unusual flow phenomena that develop on the OGVs as 

a result of the changes brought on by increasing the fan 

nozzle exit area. As expected, the stage total pressure in 

figure 11a is lower compared with the fan pressure ratio 

because of the loss in pressure across the OGVs themselves.

Again using the BL nozzle as the reference, the change in 

pressure ratio from LF to HF was from 0.9 to –6.9 percent at 

50 percent speed and 0.9 to –12.0 percent at 100 percent

speed. The stage adiabatic efficiency in figure 11b went up 

about 2.1 percent at 50 percent fan speed to 1.7 percent at 

100 percent speed, while the peak efficiency went from 

88.0 percent for the LF nozzle to 89.7 percent for the DP 

nozzle. The difference in the stage adiabatic efficiency across 

the range in nozzle area is slightly less sensitive than the fan, 

with smaller differences between nozzles for the stage 

compared to the fan. The HF nozzle performance again fell 

off at the 100 percent speed line, dropping about 2 percent

from its peak efficiency of 89.4 percent at a fan speed of 

87.5 percent, which is the same drop as seen for the fan 

efficiency with the HF nozzle.

The change in OGV performance with the change in fan 

nozzle exit area can also be expressed in terms of a loss 

function. Figures 11c and d, the total pressure and adiabatic 

efficiency losses, respectively, associated with the OGVs 

over the stage operating map are given. The total pressure 

loss across the OGVs in figure 11c, expressed as a function 

of the upstream total pressure, is defined in percent as:

(Pt,f – Pt,s)/Pt,f × 100,

In figure 11d, the loss in adiabatic efficiency across the 

OGVs is defined in percent as:

(ηf - ηs) × 100.

Figure 11c shows that the total pressure loss across the 

OGVs decreased as the nozzle area increased from the LF to 

the HF configuration. The improvement in pressure rise was 

anywhere from a minimum 0.05 percent (50% fan speed) to a 

maximum of 0.12 percent (95% fan speed) between the four 

fixed nozzle operating lines. Figure 11d shows similar results

with an increase in the adiabatic efficiency as the nozzle area 

increased. The drop in HF nozzle efficiency previously seen 

in the total pressure results can readily be seen, showing the 

drop off starting at 87.5 percent fan speed. The efficiency 

difference between nozzle configurations was shown to get 

smaller as the fan speed increased. The DP nozzle showed 

the best performance overall, since the loss in efficiency is 

the smallest at all fan speeds. This was to be expected since 

the DP nozzle achieved the design point conditions for the 

fan stage and hence the best performance. The efficiency 

increases ranged from 0.43 percent (LF to DP nozzles) at 

60 percent fan speed to 0.12 percent (LF to BL) at 

100 percent fan speed.

Figure 12 provides the spanwise radial profiles of fan 

pressure ratio (fig. 12a), temperature ratio (fig. 12b) and 

adiabatic efficiency (fig. 12c) downstream of the fan for 

several fan speeds comparing the four fan nozzle 

configurations. These figures are useful in determining what 

part of the fan blade is responsible for the change in 

performance as the nozzle exit area is changed. For the sake 

of clarity, the adiabatic efficiency plots have been separated 

into their respective fan speeds, instead of presenting all fan 

speeds on one plot. The results show that the fan loading 

increases outboard as the nozzle exit area is reduced, 

especially the outer 40 percent of the fan blade. In the 

pressure and temperature ratio plots (figs. 12a and b), the fall 

off in performance over the outer 40 percent span of the 

blade for the HF nozzle at 100 percent fan speed can readily 

be seen, especially at 70 percent span where a significant 

drop causes a bend or kink in the radial distributions. The 

adiabatic efficiency plot shows the drop-off in fan 

performance over the outer portion of the fan blade at 

87.5 percent fan speed when the fan flow first becomes 

transonic, and a recovery in performance in this region at 

100 percent fan speed once the tip flow has stabilized. The 

HF nozzle efficiency data, however, does not follow this 

trend. It shows a dramatic loss in performance outboard of 

about 50 percent span. This is especially the case at 

70 percent span where the local efficiency drops 5.5 percent

from 87.5 percent speed to 100 percent speed. This is 

possible evidence of a strong shock/flow interaction 

phenomenon occurring near 70 percent of the blade span.

The data also suggest that the performance outboard of 

70 percent span accounts for the difference in efficiency for 

the LF and BL nozzles compared with the DP nozzle—the 

DP nozzle performance is 4 to 5 percent better than the BL or 

LF nozzle performance at about 83 percent blade span at 

100 percent fan speed.

In figure 13, swirl angle data obtained downstream of the 

fan from a traversing survey probe are shown for each of the 
NASA/TM—2013-214029 9

four fan nozzle configurations at several fan speeds. These 

results demonstrate the difference in the capacity of the blade 

to do work, as well as a measure of blade loading radially 

between the four nozzles at each fan speed. Typically, as the 

blade loading goes up, the swirl angle increases. However, 

this is not necessarily true at the hub and the tip. Here, the 

plots demonstrate that the blade work capacity is reduced at 

the hub and the tip, and is most effective near the mid region.

The plots show the shape of the loading distribution remains 

fairly constant between the fan nozzles at 61.7 and 75 percent

fan speed, and that the fan performs best with the HF nozzle 

and worst with the LF nozzle. That trend is verified by the 

fan adiabatic efficiency plots shown in figure 12c. The 

loading begins to fall off near the hub at 10 percent span and 

near the tip at 85 percent span, with the peak loading 

occurring near 85 percent span. As the fan speed increases to 

87.5 percent, the BL and LF nozzles develop small variations 

in the loading distribution, as evidenced by the wiggles near 

60 and 85 percent span. These wiggles indicate loading 

changes and coincide with sharp changes in the efficiency 

profiles, appearing as losses, in figure 12. Also, the radial 

location of the fall off point in swirl seems to be moving 

inboard as nozzle area is decreased. For the LF nozzle, the 

peak loading moved from about 79 percent to about 

75 percent span. At 100 percent fan speed, the results show 

that the peak loading on the fan blades is spread over a larger 

span of the blades for the LF and BL nozzles compared to the 

DP and HF nozzles. The difference in the swirl angle 

between the BL and LF nozzle cases compared to the DP and 

HF nozzles at the peak loading point on the blades also 

seems to be increasing at the higher speed. This indicates that

the fan is not performing as well outboard for the BL and LF 

nozzles compared to the DP and HF nozzles. Finally, there is 

some indication in the swirl results from about 60 to 

70 percent blade span of the fan performance loss for the HF 

nozzle at 100 percent speed, with a increase in the swirl angle 

at that point. The adiabatic efficiency results in figure 12c

confirm this result, showing a large change in fan efficiency 

for the HF nozzle at the same blade span location.

The thrust generated by the fan and OGVs are important 

parameters for farfield acoustics since differences in 

acoustics associated with different hardware configurations 

can only be reasonably compared on an equal thrust basis.

Figure 14 shows the corrected fan thrust and corrected stage 

total thrust results obtained for BL, DP and HF nozzle 

configurations across the fan speed range on their fixed 

operating lines. The stage or total thrust is the combined 

forces from the fan and the OGV assembly. Thrust data were 

not obtained with the LF nozzle, so that configuration is not 

shown. The fan thrust results show that as the nozzle exit 

area increases, the fan thrust decreases. This is expected 

since the pressure ratio for the fan decreases as well with 

increasing nozzle area. The significant drop in performance 

for the HF nozzle at 100 percent fan speed can be seen in the 

plot. The results obtained for the corrected total stage thrust, 

which includes the fan and the OGVs assembly forces, are 

also shown in figure 14. Interestingly, the results show that as 

the fan nozzle exit area increases, the total thrust increases as 

well. The exception again is the 100 percent speed condition 

for the HF nozzle.

In figure 15, plots of the percent differences in corrected 

fan and total or stage thrust with differences in nozzle exit 

area are shown for selected fan speeds. For this figure, the 

BL nozzle results were used as the zero reference point. The 

thrust results presented in this form clearly demonstrate the 

increase in corrected total thrust achieved by the OGVs 

across almost the entire fan speed range associated with 

increasing the nozzle area. It was shown in figure 14 that the 

fan thrust decreases with increasing nozzle area, and that 

result is shown in this figure. The increase in total thrust 

shown in figure 15 must therefore come from an increase in 

thrust produced by the OGVs. The results also indicate that 

as the fan speed increases, the total thrust advantage 

decreases slightly. The loss in fan thrust from figure 15 

ranges between –0.2 percent at low speeds to –3.3 percent at 

high speeds for the DP nozzle, and between –3.8 and 

–7.7 percent for the HF nozzle, with the exception of the 

takeoff fan speed that shows a drop of –10.2 percent in 

thrust. The increase in total thrust ranges from around 2

percent at higher fan speeds to about 3.5 percent at lower fan 

speeds compared to the BL nozzle, depending on the fan 

speed. The results also show that the increase in total thrust 

produced is almost constant, or drops slightly, as the nozzle 

area increases from the DP nozzle to the HF nozzle, again 

except for 100 percent fan speed. The large loss in fan thrust 

that occurs for the HF nozzle at 100 percent fan speed

produces an overall loss in total thrust compared to the BL 

nozzle, but that loss is significantly reduced by the increase 

in thrust from the OGVs. The reason for the increase in total

thrust is most likely due to the increase in the lift force 

generated by the vanes in the direction of the thrust vector, as 

a result of the increase in the axial velocity that accompanies 

the increase in weight flow as the fan nozzle exit area 

increases. 

Farfield Acoustics

Figure 16 shows the overall sound power level (OAPWL)

as a function of stage thrust for three fixed area nozzle

configurations. The OAPWL was calculated from the 59 Hz 

bandwidth (BW) spectra over a 1 to 50K frequency range for 

sideline emission angles from 25 to 158°

. The results show a 

significant noise decrease associated with increasing nozzle 

area and weight flow—especially at lower fan speeds.

Figure 17 shows the change in OAPWL plotted for the DP 

and HF nozzles relative to the BL nozzle. These delta 

OAPWL were measured at constant thrust levels interpolated 

from the curves of figure 16. Increasing the nozzle area 

5.4 percent and the weight flow by 5 percent with the DP 

nozzle resulted in a 2 dB noise reduction at rotor speeds up to 

cutback (87.5% design speed), and around a 1 dB reduction 

at higher rotor speeds. Further increasing the nozzle area to 
NASA/TM—2013-214029 10

10.9 percent and the weight flow to 7.5 percent of BL design 

values gave noise reductions of about 3 dB relative to the BL 

nozzle at lower rotor speeds (near approach at 61.7% design 

speed) and a 2 dB noise reduction at intermediate rotor 

speeds. However, the more open HF nozzle resulted in higher 

noise levels at the highest rotor speeds. This is the likely 

result of the degraded rotor performance with the HF nozzle 

shown earlier in the aerodynamic performance section near 

the 100 percent fan design speed.

Noise reductions associated with increased nozzle flow 

extend over a significant frequency range. Figure 18 shows

the sound power level spectra (PWL) results obtained at the 

three fan stage acoustic rating conditions—approach, cutback 

and takeoff, corresponding to 61.7, 87.5, and 100 percent

corrected fan speed. For the 61.7 percent fan speed presented 

in figure 18a, the blade/vane ratio numbers for this fan stage 

design resulted in the fundamental rotor-stator interaction 

tone (ref. 30), or blade passing frequency (BPF), being 

essentially eliminated from the spectra, or cut off. This is true

at fan speeds below where the relative velocity on the fan 

blade is below transonic, or for this fan at 87.5 percent speed 

(cutback). However the 2nd BPF tone is visible in the spectra

and shows that there is a slight increase in tone level as the 

nozzle area and hence weight flow are increased. The 2nd

BPF is not present at the other two fan speeds shown in the 

figure because beyond 8 kHz there is a significant artificial 

roll-off in the level of the data as a result of the bandwidth 

packets selected for data reduction to get higher data 

resolution for these plots. The reason for the increase in tone 

level between nozzle configurations may be that increasing 

the weight flow as the nozzle area increases also increased 

the axial velocity component of the flow. As a result, the fan 

relative velocity and fan tip speed also increase, and since 

tone level is partly a function of fan tip speed, the tone level 

increases with the larger nozzles. A significant change in the 

broadband noise levels between nozzles can be seen in the 

figure across the entire frequency range. The broadband 

noise levels are up to 3 dB lower for the DP nozzle and 6 dB 

lower for the HF nozzle relative to the noise levels observed 

for the BL nozzle. A possible explanation is the reduced 

loading on the fan blades in terms of lower fan pressure ratio

(Figs. 10b and 12a), reduced fan swirl (fig. 13), and lower 

rotor thrust (fig. 14) that was shown in the aerodynamic 

performance results section.

At 87.5 percent corrected fan speed, there is a significant 

increase in the BPF tone level with increasing nozzle area 

(fig. 18b). This may be explained by the previous argument 

of increasing rotor relative velocity with increasing weight 

flow at larger nozzle areas. The higher fan tip velocity for the 

HF and DP nozzles is generating stronger shocks than for the 

BL case. As the flow tip reaches sonic flow conditions, rotoralone noise begins to dominate the rotor tone levels. In 

addition, the onset of supersonic flow on the fan initiates the 

generation of Multiple Pure Tones (MPTs). MPTs are tones 

which occur in the acoustic spectra at multiples of the rotor 

once per revolution frequency. They are thought to result 

from blade-to-blade differences in the shocks which 

propagate upstream of the rotor when the rotor is operated at 

transonic tip. The MPTs become more evident with 

increasing nozzle area and weight flow at this cutback fan 

speed, especially for the higher flow nozzles (This earlier 

onset of MPTs with increased nozzle flow is consistent with 

the OAPWL results shown in fig. 17). Apparently, opening 

the nozzle causes the shocks on the blades to become 

stronger and to extend over a larger spanwise extent of the 

blade. This, in turn, results in more shock associated noise.

The spectra of figure 18b show a modest reduction in 

broadband noise of about 2 dB with both larger area nozzles

for this fan speed. The noise “hump” at about 1⁄2 BPF for the 

BL nozzle is unexplained.

At takeoff conditions at 100 percent corrected fan speed, 

the MPT generation is well established (fig. 18c). At this 

speed the shocks extend over a large spanwise extent of the 

blades. There is a small increase in the BPF tone level with 

increasing nozzle flow, but the increase is smaller between 

the HF and DP nozzles. This relatively small increase may be 

associated with the location of the shocks on the blades. At 

these high flow conditions it is likely that the shocks in the 

tip region have been pushed downstream to the point where 

they bend around the leading edge of the adjacent blades 

(confirmed by LDV measurements presented in ref. 15).

Once the shocks begin to bend around the adjacent blades, 

further increases in flow velocity cause the shocks to bend 

even more. This increased bending moves the shocks 

downstream toward the rotor face and away from the inlet 

throat. This is important since it is the amplitude of the 

disturbance created by the shocks at the inlet throat which 

dictates how much shock associated noise is radiated to the 

farfield. Consequently, the increase in flow velocity which 

results from opening the nozzle at the takeoff condition can 

be expected to lead to only modest increases (or perhaps even 

reductions) in shock associated noise (ref. 31). The 

broadband noise levels show a modest reduction at low 

frequencies with increasing nozzle area, but not nearly as 

dramatic as the previous two fan speeds. The broadband 

noise is also now dominated by the supersonic flow over the 

fan blades and the shock-fan blade interaction noise, so the 

modest changes in fan blade loading, which appears to be the 

main driver for broadband noise as previously shown in the 

figures 18a and b, are not significant.

Finally, figure 19 provides a summary plot of the change 

in OASPL as a function of fan nozzle exit area change for 

various fan speeds tested from 50 to 100 percent. The Low 

Flow nozzle is not represented in the figure since no acoustic 

data was obtained with that nozzle. The figure provides a 

summary of the acoustic benefit which results from

increasing the fan nozzle exit area has on the noise produced 

by the fan module. It can be seen that for most operating 

conditions a significant decrease in the OASPL is possible, 

the exception being the high speed conditions at cutback and 

takeoff for the HF nozzle. However, even for those 

conditions a modest decrease in noise was seen.
NASA/TM—2013-214029 11

Laser Doppler Velocimetry

The LDV data were obtained in order to determine how 

the wake flow generated by the fan changes as the nozzle 

area changes. These wake data can be used to explain some 

of the observations made above regarding how the acoustic 

spectra change as the nozzle area and weight flow are

increased.

Figures 20 through 22 show how the rotor wake flow 

varies with changes in nozzle area as measured at the 

approach, cutback and takeoff fan operating conditions, 

respectively. Parts a and b of these figures show that the 

mean axial velocities increase and the mean tangential 

velocities decrease in the rotor wake as nozzle area increases.

As a result, swirl angle also decreases, as was shown in 

figure 13. As was suggested then, the decreased swirl 

suggests that the loading on the rotor blades decreases as 

nozzle area increases. In general, this decreased blade 

loading should lead to a decrease in the amount of turbulence 

generated by each blade. The plots shown in parts c) and d) 

of each figure, which depict the axial and tangential 

turbulence components, confirm that this was the case—the 

measured turbulence level in the rotor wake decreased as 

nozzle area increased. This decreased rotor wake turbulence 

will result in the decreased levels of rotor/stator interaction 

broadband noise. This was evident in the acoustic spectra 

plots of figure 18a and b. These plots showed a reduction in 

broadband noise with increasing nozzle area, as discussed 

earlier in the section on farfield acoustic results.

However, the broadband acoustic data presented for the 

takeoff condition in figure 18c show a different trend. These 

data are similar to the approach and cutback condition data in 

that they show a decrease in the broadband level as the 

nozzle area is increased from 0 (BL nozzle) to 5.4 percent

open (DP nozzle), but they are unlike the other data in that 

they show an increase in high frequency broadband noise as 

the nozzle area is increased further to 10.9 percent open (HF 

nozzle). The LDV flow field data can also be used to explain 

this anomaly. Figure 22 shows the variation in rotor wake 

flow with increasing nozzle area as measured at the take off 

condition. Like the data obtained at the two lower fan speeds, 

the axial velocities (part a) increase while the tangential 

velocities (part b) decrease as nozzle area increases.

Consequently, swirl angle also decreases at takeoff, as was 

shown in figure 13. The reduced swirl implies that the blade 

loading decreases with increasing nozzle area—a result that 

should lead to less turbulence generated by the blades.

However, as indicated in parts c and d of figure 22, the 

turbulence generated by the blades actually increases over 

much of the blade span as the nozzle area and weight flow 

increase between the DP nozzle and the HF nozzle.

Figures 22c and d also indicate that the outer portions of the 

blade wakes get thicker and more turbulent as the nozzle area 

increases between the DP and HF nozzles at the takeoff speed.

This increased blade wake thickness may be indicating a flow 

separation on the fan blades resulting, perhaps, from stronger 

shocks on the blades or from non-optimum inflow angles. In 

any event, this increase in blade wake turbulence seems to 

account for the increase in high frequency broadband noise 

illustrated in figure 18c for the takeoff condition. These thicker 

wakes are also consistent with the degraded aerodynamic 

performance measured at this test condition shown earlier.

This was especially evident in the fan adiabatic efficiency plots 

shown in figures 10d and 12c, and in the changes in corrected 

total thrust shown in figure 15.

The trends indicated by the LDV data also support the 

explanation presented above in the farfield acoustics result 

section earlier for the increase in BPF and MPT noise which 

occurs when the nozzle is opened at the 87.5 percent cutback 

speed. Previous reports (refs. 4 and 5) have presented LDV 

data obtained within the rotor blade passages at a radial 

location 0.4 in. inboard of the tip during a test in which the 

baseline nozzle was installed. These data show that normal 

shocks exist on the suction side of the blades when the fan is 

operating at the cutback speed. The plot presented above in 

figure 21a show that axial flow velocities increase as the area 

increases above that of the BL nozzle. The increased axial 

velocities would lead to higher relative flow velocities on the 

blades which, in turn, would be expected to lead to stronger 

passage shocks. The increased noise produced by these 

shocks is evident in the acoustic spectra of figures 18b and c,

which show that both the BPF tone and the Multiple Pure 

Tones increase as the nozzle area increases.

Summary and Conclusion

The design of effective new technologies to reduce aircraft 

propulsion noise is dependent on identifying and 

understanding the noise sources and noise generation 

mechanisms in the modern turbofan engine, as well as 

determining their contribution to the overall aircraft noise 

signature. Therefore, a comprehensive aeroacoustic wind 

tunnel test program was conducted as part of the NASA 

Quiet Aircraft Technology program called the Fan 

Broadband Source Diagnostic Test. The test was performed 

in the anechoic NASA Glenn 9- by 15-Foot Low Speed Wind 

Tunnel using a 1/5 scale model turbofan simulator that is 

representative of a current generation, medium pressure ratio 

high bypass turbofan engine. The investigation focused on 

the simulated bypass section of the turbofan engine. The 

technical objectives of the test were: 1) to identify the noise 

sources within the model and their contribution to the overall 

noise level; 2) to investigate several component design 

technologies by evaluating their impact on the aerodynamic 

and acoustic performance; and 3) to conduct detailed flow 

diagnostics within the research model to help in 

understanding the physics of the flowfield.

Details were presented in this report on the aerodynamic 

performance, farfield acoustics and fan wake flow diagnostic 

results obtained during an investigation of the effect of the 

bypass nozzle exit area on the bypass stage performance, 
NASA/TM—2013-214029 12

specifically the fan and outlet guide vanes (OGVs). Four 

different fixed-area bypass nozzles were investigated in 

combination with a baseline fan and set of OGVs—Low 

Flow (LF), Baseline (BL), Design Point (DP) and High Flow

(HF). They represented fixed engine operating lines 

encompassing the operating envelope of the turbofan engine 

from near stall to cruise, with a total change in area from the 

smallest to the largest nozzle of 12.9 percent. The BL nozzle 

was selected as the reference nozzle area. Its area was 

2 percent larger than the LF nozzle, 5.4 percent smaller than 

the DP nozzle and 10.9 percent smaller than the HF nozzle. 

The results demonstrate that there are significant changes 

in aerodynamic performance and farfield acoustics as the 

nozzle area is increased:

1) Fan and stage pressure and temperature ratios decrease 

as nozzle area increases. The maximum loss in total pressure 

from LF to HF was 12.5 percent at takeoff condition (100%

fan design speed). However, the HF nozzle performance was 

seriously degraded at the 100 percent speed point most likely 

due to shock losses on the blades. At 95 percent speed, just 

below takeoff, the overall pressure loss was only 9.7 percent.

The maximum temperature loss between LF and HF 

remained fairly constant with fan speed, between 10.7 and 

11.3 percent. For the stage, the maximum loss in pressure 

was 12.9 percent at 100 percent speed and 10.1 percent at 

95 percent fan speed.

2) Fan and stage adiabatic efficiency increase as the nozzle 

area increase, the fan and stage efficiency increased as well.

The overall change in fan efficiency from LF to HF nozzle 

varied slightly with fan speed. The overall change was 

generally from 2.5 to 3 percent, except at the 100 percent

point, where the maximum difference was only 1.1 percent.

For the stage adiabatic efficiency, the overall change between 

the LF and HF nozzles was between 2.4 and 3.1 percent up to 

87.5 percent fan speed. At higher fan speeds, the increase in 

stage efficiency associated with larger nozzle area still occur 

but was smaller, a 1.8 percent increase at 100 percent fan 

speed. The smaller increase is a result of increased losses on 

the OGVs at the higher weight flow and axial velocity.

3) Radial profile results show that the HF nozzle provided 

the lightest blade loading and the highest adiabatic 

efficiency. The shock induced losses on the blade at 100

percent fan speed for the HF nozzle could easily be seen in 

the efficiency profile near 70 percent blade span. The swirl 

angle survey results showed the drop in blade loading near 70

percent span at 100 percent fan speed, the location which 

showed the large drop in local efficiency and the area where 

flow separation on the fan blade due shock interactions might 

have occurred. The swirl angle results also showed that the 

BL and LF nozzles produced highly loaded blades above 70

percent span which caused the fan efficiency to drop near the 

tip.

4) Corrected fan thrust decreases with increasing nozzle 

area, but corrected stage thrust—the combination of fan and 

OGV thrust—increases with increasing nozzle area up to area 

increases of about 5 percent. Beyond that, as the nozzle exit 

area continues to increase, the stage thrust advantage levels 

off and remains constant or even drops at very low or very 

high fan speeds. Compared to the BL nozzle, the overall fan 

thrust loss for the DP nozzle was from 0.3 percent

at approach (61.7% fan speed) to 2.3 percent at takeoff 

(100% fan speed), while the fan thrust loss for the HF nozzle 

ranged from about 3.8 percent at approach to 6.7 percent at 

95 percent fan speed. A jump in losses to 10.3 percent

occurred at 100 percent fan speed as the shock losses on the 

fan blade dramatically reduce the blade performance. The 

stage total thrust increases compared with the BL nozzle 

reference from 3.6 percent at approach to 2.1 percent at 

takeoff for the DP nozzle. As the nozzle area increases for 

the HF nozzle, the gain in total thrust remains fairly constant 

or decreases slightly by about 0.2 percent, except at 

100 percent fan speed. Here, the stage thrust drops compared 

with the BL nozzle stage thrust, losing 1.9 percent in thrust.

5) As the nozzle area increases, the overall sound power 

level decreases anywhere from 3.5 dB at approach speed 

(61.7% fan speed) to 1 dB at takeoff (100% fan speed). The 

HF nozzle showed a slight 1dB advantage over the DP nozzle 

at lower speeds (up to about 75% fan speed), but the 

advantage disappears at the higher speeds, except around 

90 percent speed where the HF again had the 1 dB advantage.

At 95 percent speed and above, the HF nozzle was noisier 

than the DP or BL nozzles due, presumably, to shockinduced flow separation and losses. The reductions in overall 

noise level associated with increasing the nozzle area were 

primarily a result of decreased broadband noise. The fan tone 

levels actually increased with increasing nozzle area, most 

likely due to the higher relative speed on the fan blades 

because of the increase in weight flow and hence axial 

velocity as the nozzle area increased.

6) Laser Doppler velocimeter results show that mean axial 

velocities increase and mean tangential velocities decrease 

with increasing nozzle area at a given fan speed.

Consequently, the swirl in the wake flow also decreases—a 

result which indicates that the fan blade loading decreases 

with increasing nozzle area. At the approach and cutback 

speeds, wake turbulence levels decrease and the blade wakes 

get thinner as the nozzle area is increased. These trends 

explain the drop in broadband noise level measured as the 

nozzle area increased. At 100 percent speed, the results 

showed an increase in wake turbulence when the HF nozzle 

was installed. This is thought to result from shock losses in 

the outboard region of the fan blade. Again, this correlates 

with the acoustic results which showed the broadband level 

increasing at high frequencies for the HF nozzle at takeoff.

In summary, the results provided in this report suggest that 

a variable area bypass exhaust nozzle for a typical turbofan 

engine may be an effective way to further decrease engine 

fan stage noise and possibly realize a concurrent slight thrust 

increase. Turbofan engine bypass exhaust nozzles are 

normally sized for maximum performance at the portion of 

the aircraft flight profile where most of the flight time is 

spent–typically at the cruise condition. Increasing the nozzle 
NASA/TM—2013-214029 13

flow within the envelope defined by desirable engine 

performance reduced the fan stage noise in this scale model 

test. Thus, it may be desirable to employ a variable area 

engine bypass exhaust nozzle as a technique to reduce fan 

stage noise levels at all rotor operating speeds. Even the 

addition of a limited position variable area bypass nozzle, in 

order to reduce mechanical complexity and weight, might be 

an effective retrofit to existing turbofan engines to control 

fan stage noise and realize additional noise reduction without 

sacrificing aerodynamic performance.

References

1. Groeneweg, J.F., and Rice, E.J., “Aircraft Turbofan Noise,”

vol. 109, Journal of Turbomachinery, January 1987.

2. Groeneweg, J.F., “Fan Noise Research at NASA,” NASA TM–

106512, 1994.

3. Envia, E., “Fan Noise Reduction: An Overview,” NASA/TM—

2001-210699, AIAA–2001–0661, 2001.

4. Bridges, J., Envia, E., and Huff, D., “Recent Developments in 

U.S. Engine Noise Reduction Research,” NASA/TM—2001-

211083, ISABE–2001–1017, 2001.

5. Dittmar, J.H. and Elliott, D.M., and Bock, L.A., “Some 

Acoustic Results From the Pratt & Whitney Advanced Ducted 

Propulsor—Fan 1,” NASA/TM—1999-209049; March 1999.

6. Dittmar, J.H. and Hughes, C.E., Bock, L.A., and Hall, D.G.,

“Cruise Noise Measurements of a Scale Model Advanced 

Ducted Propulsor,” 15th AIAA Aeroacoustics Conference, 

Long Beach, CA; AIAA–93–4400, October 25–27, 1993.

7. Envia, E., and Nallasamy, M., “Design Selection and Analysis 

of a Swept and Leaned Stator Concept,” NASA/TM—1998-

208662, December 1998.

8. Hoyniak, D. and Fleeter, S., “Effect of Aerodynamic Detuning 

on Supersonic Rotor Discrete Frequency Noise Generation,”

NOISE-CON 88, Purdue University, West Lafayette, IN; June 

20–22, 1988

9. Woodward, R.P., Elliott, D.M., Hughes, C.E., and Berton, J.J.,

"Benefits of Swept and Leaned Stators for Fan Noise 

Reduction,” NASA/TM—1998-208661; AIAA–99–0479; 

November 1998.

10. Podboy, G.G., “Further Analysis of Fan Wake Data Obtained 

Downstream of the Allison Low Noise Fan and the Pratt & 

Whitney ADP “Fan 1’ Models,” AST Engine Noise Workshop, 

vol. IV, pp. 131−160, April 1998.

11. Hughes, C.E., “Aerodynamic Performance of Scale-Model 

Turbofan Outlet Guide Vanes Designed for Low Noise,” 

AIAA–2002–0374, January 2002.

12. Podboy, G.G., Krupar, M.J., Helland, S.M., and Hughes, C.E., 

“Steady and Unsteady Flow Field Measurements Within a 

NASA 22 Inch Fan Model,” AIAA–2002–1033, January 2002.

13. Woodward, R.P., Hughes, C.E., Jeracki, R.J., and Miller, C.J., 

“Fan Noise Source Diagnostic Test—Far-field Acoustic 

Results,” AIAA–2002–2427, June, 2002. 

14. Hughes, C.E., Jeracki, R.J., and Miller, C.J., “Fan Noise 

Source Diagnostic Test—Rotor Alone Aerodynamic 

Performance Results,” AIAA–2002–2426, June 2002.

15. Podboy, G.G., Krupar, M.J., Hughes, C.E., and Woodward, 

R.P., “Fan Source Diagnostic Test—LDV Measured Flow 

Field Results,” AIAA–2002–2431, June 2002.

16. Heidelberg, L.J., “Fan Noise Source Diagnostic Test—Tone 

Model Structure Results,” AIAA–2002–2428, June 2002.

17. Woodward, R.P., Lucas, J.G., and Stakolilch, E.G., “Acoustic 

and Aerodynamic Performance of a 1.83-Meter- (6-Ft-) 

Diameter 1.2-Pressure-Ratio Fan (QF-6),” NASA TN D–7809, 

December, 1974.

18. Woodward, R.P., Lucas, J.G., and Balombin, J.G., “Acoustic 

and Aerodynamic Performance of a 1.5-Pressure-Ratio, 1.83-

Meter (6 Ft.-) Diameter Fan Stage for Turbofan Engines (QF2),” NASA TM X–3521, April, 1977.

19. Woodward, R.P., and Glaser, F.W., “Effect of Inflow Control 

on Inlet Noise of a Cut-on Fan,” AIAA Journal, vol. 19, no. 3,

March 1981, pp. 387–392.

20. Gliebe, P.R., “The Effect of Throttling on Forward Radiated 

Fan Noise,” AIAA–79–0640, March, 1979.

21. Ginder, R.B. and Newby, D.R., “An Improved Correlation for 

the Broadband Noise of High-Speed Fans,” AIAA Journal of 

Aircraft, vol. 14, no. 9, September, 1977, pp. 844–849.

22. Yuska, J.A., Diedrich, J.H., and Nestor, C., “Lewis 9- by 15-

Foot V/STOL Wind Tunnel,” NASA TM X–2305, 1971.

23. Arrington, A.E., and Gonsalez, J.C., “Flow Quality 

Improvements in the NASA Lewis Research Center 9- by 15-

Foot Low Speed Wind Tunnel,” NASA CR–195439, 1995.

24. Dahl, M.D., and Woodward, R.P., “Comparison Between 

Design and Installed Acoustic Characteristics of the NASA 

Lewis 9- by 15-Foot Low Speed Wind Tunnel Acoustic 

Treatment,” NASA TP–2996, April 1990.

25. Dahl, M.D., and Woodward, R.P., “Background Noise Levels 

Measured I the NASA Lewis 9- by 15-Foot Low Speed Wind 

Tunnel,” NASA TP–3274, November 1992.

26. Woodward, R.P., and Dittmar, J.H., “Background Noise Levels 

Measured in the NASA Lewis 9- by 15-Foot Low-Speed Wind 

Tunnel,” NASA TM–106817, AIAA–95–0720, January 1995.

27. Balan, C. and Hoff, G.E., “Propulsion Simulator for High 

Bypass Turbofan Performance Evaluation,” SAE 931410, 

1993.

28. Jeracki, R.J., “Model Engine Performance Measurement from 

Force Balance Instrumentation,” NASA/TM—1998-208486, 

AIAA–98–3112, July 1998.

29. Chestnutt, D., “Flight Effects of Fan Noise,” NASA CP–2242, 

January 1982.

30. Tyler, J.M. and Sofrin, T.G., “Axial Flow Compressor Noise 

Studies,” SAE Trans, vol. 70, 1962, pp. 309–332.

31. Podboy, G.G. and Krupar, M.K., “Blade-to-Blade Variations in 

Shocks Upstream of Both a Forward-Swept and an Aft-Swept 

Fan,” NASA/TM—2005-213445, 2005.
NASA/TM—2013-214029 14

TABLE 1.—FAN DESIGN PARAMETERS

No. of Blades 22

Tip Diameter, in. 22

Inlet Radius Ratio 0.30

Corrected Design Speed, rpm 12,657

Design Tip Speed, ft/s 1,215

Corrected Weight Flow, lbm/s 100.5

Specific Flow, lbm/s-ft2 41.8

Stage Pressure Ratio 1.47

TABLE 2.—OUTLET GUIDE VANE DESIGN PARAMETERS

Span Location

No. Vanes 54

Aft Sweep, deg 0

Aspect Ratio Pitchline 3.51

Chord, in Pitchline 1.57

Solidity Hub 2.25

Pitchline 1.52

Tip 1.23

Stagger, deg 1,2 Hub 12.56

Pitchline 10.29

Tip 10.65

Vane Camber, deg Hub 38.40

Pitchline 34.56

Tip 40.49

tmax/c Hub 0.0707

Pitchline 0.0702

Tip 0.0698

1 Defined from axial plane; positive angle in direction of fan rotation.

2 Positive angle in opposite direction of fan rotation for OGVs.

TABLE 3.—FAN EXIT NOZZLE DESIGN PARAMETERS

Nozzle Weight Flow, 

lbm/sec

Diameter, 

in

Area, 

in2

Baseline 97.2 21.340 217.38

Low Flow 96.1 21.210 213.04

Design Point 101.4 21.684 229.01

High Flow 102.7 22.034 241.02
NASA/TM—2013-214029 15

Figure 1.—NASA Glenn Research Center 8- by 6-Foot Supersonic Wind Tunnel/

9- by 15-Foot Low Speed Wind Tunnel Complex.

Screen and honeycomb

flow conditioners

8x6 test section

0 ≤ M ≤ 0.1 and

0.25 ≤ M ≤ 2.0

Turn 1 Flow High speed diffuser

Screen and honeycomb

flow conditioners

Cooler

Turn 2 Acoustic muffler

37.856°

Grid North

N

Turn 3

Air dryer building

with 8 large fans

Flow

Low speed diffuser

9x15 test section

0 ≤ M ≤ 0.2

Doors 4 and 5

Doors 

1 and 2

Drive motor building

housing three 29000 hp

electric motors

Seven stage

axial compressor

Balance chamber

Flexible

wall

nozzle

Door 3

Shock

doors

 
NASA/TM—2013-214029 16

a) Aerodynamic Performance configuration

b) Farfield Acoustics configuration

Figure 2.—Top views showing test hardware locations in Glenn 9- by 15-Foot Low Speed Wind Tunnel.

Fan Stacking Axis

MS 170.000

UHB Drive Rig

Fan Model Installed with 

Bellmouth Inlet and 

Variable Fan Exit Nozzle 

(VFEN)

End of Acoustic Boxes End of Test Section

Traversing Microphone 

Mechanism Enclosure

Test Section Centerline

Cruciform Rake

Start of Test Section

Start of Acoustic Boxes

Flow

335.9”

344.0”

178.7”

89.25”

30.0”

59.4”

4.0” 10.0”

21.5” 16.0”

176.0”

121.0” 67.0”

207.55”

27.0”

Instrumentation 

Egress Tubes

Dimensions in inches

Fan Stacking Axis

MS 170.000

UHB Drive Rig

Fan Model Installed with 

Bellmouth Inlet and 

Variable Fan Exit Nozzle 

(VFEN)

End of Acoustic Boxes End of Test Section

Traversing Microphone 

Mechanism Enclosure

Test Section Centerline

Cruciform Rake

Start of Test Section

Start of Acoustic Boxes

Flow

335.9”

344.0”

178.7”

89.25”

30.0”

59.4”

4.0” 10.0”

21.5” 16.0”

176.0”

121.0” 67.0”

207.55”

27.0”

Instrumentation 

Egress Tubes

Dimensions in inches

Fan Stacking Axis

Model STA 170.000

UHB Drive Rig

Traversing Acoustic Microphone 

Flow

Test Section 

Centerline

335.9”

344.0”

176.0”

207.55”
NASA/TM—2013-214029 17

Figure 3.—NASA Glenn Research Center Ultra High Bypass (UHB) Drive Rig propulsion simulator.

Figure 4.—Schematic diagram of the Fan Model and Baseline OGVs installed on the UHB Drive Rig.

 

Operating Capabilities

Maximum Power 5,050 HP

Maximum Speed 16,850 RPM

Turbine Plenum Conditions 

@ Max Operation

Pressure 230 psia

Temperature 500°F

Mass Flow 33.6 lbs/sec

Air Turbine Drive

Drive Rig 

Support Strut

Fan Module Location

Fan 

Flight-Type Nacelle

Fixed Area Flight-Type Nozzle

Baseline 

Flight OGVs -Type Inlet UHB Drive Rig

Rotating Balance 

Static Force Balance

Fan 

Flight-Type Nacelle

Fixed Area Flight-Type Nozzle

Baseline 

Flight OGVs -Type Inlet UHB Drive Rig

Rotating Balance 

Static Force Balance
NASA/TM—2013-214029 18

Figure 5.—Schematic diagram of fan module hardware and variation in fan nozzle exit area.

10.5

11.0

11.5

12.0

186 187 188 189 190 191 192 193

Model Station, in

Radius, in

0

5

10

15

20

155 160 165 170 175 180 185 190 195 200 205 210 215 220 225

Model Station, in

Radius, in

Flight-type Nacelle

Fan Nozzle 

Afterbody

Fan Stacking Axis

Stator Stacking 

A i

 Low Flow

 Baseline

 Design Point

 High Flow

Nozzle Config

10.5

11.0

11.5

12.0

186 187 188 189 190 191 192 193

Model Station, in

Radius, in

0

5

10

15

20

155 160 165 170 175 180 185 190 195 200 205 210 215 220 225

Model Station, in

Radius, in

Flight-type Nacelle

Fan Nozzle 

Afterbody

Fan Stacking Axis

Stator Stacking 

A i

 Low Flow

 Baseline

 Design Point

 High Flow

Nozzle Config
NASA/TM—2013-214029 19

a) Fan module with bellmouth inlet and variable area fan exit nozzle

b) Close-up view of variable fan exit nozzle hardware arrangement

Figure 6.—Photographs of fan module installed in the Glenn 9- by 15-Foot

Low Speed Wind Tunnel in the aerodynamic performance configuration.
NASA/TM—2013-214029 20

a) LDV system hardware installed next to the research fan model in the wind tunnel

b) LDV system components with optics protective shield removed

Figure 7.—LDV system components and installation setup with 

the fan module in Glenn 9- by 15-Foot Low Speed Wind Tunnel.

Cylindrical

shield 

covering

optics

Window

Receiving optics

Transmitting optics Fiber

optic

cables
NASA/TM—2013-214029 21

.

a) Fan module with flight-type inlet, nacelle and fixed area fan nozzle

b) View of acoustic microphone orientation in the wind tunnel

Figure 8.—Photographs of fan module installed in the Glenn 9- by 15-Foot 

Low Speed Wind Tunnel in the farfield acoustic configuration.

Fixed 

microphones

Traversing 

microphone

UHB Drive Rig 

air turbine

Air supply and 

support strut
NASA/TM—2013-214029 22

Figure 9.—Schematic diagram of the LDV fan wake measurement 

plane axial location within the fan module.

Flow

Bellmouth Inlet

Fan

Swept

OGV

 LDV Measurement Plane

 (Baseline OGV leading edge)
NASA/TM—2013-214029 23

a) Corrected weight flow map on the fixed area nozzle operating lines

b) Total pressure ratio map

Figure 10.—Fan performance maps (continued).

40

50

60

70

80

90

100

110

40 50 60 70 80 90 100 110

Corrected Fan Design Speed, %

Corrected Weight Flow, lbm/sec

 LF

 BL

 DP

 HF

Nozzle

1.00

1.10

1.20

1.30

1.40

1.50

1.60

40 50 60 70 80 90 100 110

Fan Pressure Ratio, Pt,f/Pt,o

 50%

 60%

 70%

 77.5%

 87.5%

 95%

 100%

 LF Op Line

 BL Op Line

 DP Op Line

 HF Op Line

 High Stress Line

Corrected Weight Flow, ωf,c , lbm/sec
NASA/TM—2013-214029 24

c) Total temperature ratio map

d) Adiabatic efficiency map

Figure 10.—Fan performance maps (concluded).

1.00

1.02

1.04

1.06

1.08

1.10

1.12

1.14

1.16

40 50 60 70 80 90 100 110

Fan Temperature Ratio, Tt,f/Tt,o

 50%

 60%

 70%

 77.5%

 87.5%

 95%

 100%

 LF Op Line

 BL Op Line

 DP Op Line

 HF Op Line

 High Stress Line

Corrected Weight Flow, ωf,c , lbm/sec

80

82

84

86

88

90

92

94

96

98

100

40 50 60 70 80 90 100 110

 50%

 60%

 70%

 77.5%

 87.5%

 95%

 100%

 LF Op Line

 BL Op Line

 DP Op Line

 HF Op Line

Fan Adiabatic Efficiency, ηf, %

Corrected Weight Flow, ωf,c , lbm/sec
NASA/TM—2013-214029 25

a) Total pressure ratio map

b) Adiabatic efficiency map

Figure 11.—Stage performance maps (continued).

1.00

1.10

1.20

1.30

1.40

1.50

1.60

40 50 60 70 80 90 100 110

Stage Pressure Ratio, Pt,s/Pt,o

 50%

 60%

 70%

 77.5%

 87.5%

 95%

 100%

 LF Op Line

 BL Op Line

 DP Op Line

 HF Op Line

Corrected Weight Flow, ωf,c , lbm/sec

76

78

80

82

84

86

88

90

92

94

96

40 50 60 70 80 90 100 110

 50%

 60%

 70%

 77.5%

 87.5%

 95%

 100%

 LF Op Line

 BL Op Line

 DP Op Line

 HF Op Line

Corrected Weight Flow, ωf,c , lbm/sec

Stage Adiabatic Efficiency, ηf, %
NASA/TM—2013-214029 26

c) Total pressure loss across the OGVs

d) Adiabatic efficiency loss across the OGVs

Figure 11.—Stage performance maps (concluded).

0.0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

1.6

1.8

2.0

2.2

2.4

40 50 60 70 80 90 100 110

OGV Total Pressure Loss Coefficient, (Pt,f-Pt,s)/Pt,f, %

50

60

70

77.5

87.5

95

100

LF Op Line

BL Op Line

DP Op Line

HF Op Line

Corrected Weight Flow, ωf,c , lbm/sec

1

2

3

4

5

6

7

8

9

10

11

12

40 50 60 70 80 90 100 110

50

60

70

77.5

87.5

95

100

LF Op Line

BL Op Line

DP Op Line

HF Op Line

Corrected Weight Flow, ωf,c , lbm/sec

OGV Adiabatic Efficiency Loss, (ηf - ηs), %
NASA/TM—2013-214029 27

a) Total pressure ratio profiles

b) Total temperature ratio profiles

Figure 12.—Comparison of fan wake radial profiles

with nozzle area at selected fan speeds (continued).

0

10

20

30

40

50

60

70

80

90

100

1.0 1.1 1.2 1.3 1.4 1.5 1.6

Fan Rake Pressure Ratio, Pt,r/Pt,o

Blade Span, %

61.7% 75% 87.5% 100%

 Nozzle 

 LF

 BL

 DP

 HF

Tip

Hub

0

10

20

30

40

50

60

70

80

90

100

1.00 1.02 1.04 1.06 1.08 1.10 1.12 1.14 1.16 1.18

Fan Rake Temperature Ratio, Tt,r/Tt,o

Blade Span, %

61.7% 75% 87.5% 100%

Tip

Hub

 Nozzle 

 LF

 BL

 DP

 HF
NASA/TM—2013-214029 28

c) Adiabatic efficiency profiles

Figure 12.—Comparison of fan wake radial profiles with

nozzle area at selected fan speeds (concluded).

Figure 13.—Comparison of fan wake radial swirl angle 

profiles with nozzle area at selected fan speeds.

0

10

20

30

40

50

60

70

80

90

100

64 68 72 76 80 84 88 92 96 100

Blade Span, %

Fan Rake Adiabatic Efficiency, ηr, %

Tip

Hub

 Fan Speed

 61.7%

 75%

 87.5%

 100%

 Nozzle 

 LF

 BL

 DP

 HF

0

10

20

30

40

50

60

70

80

90

100

24 26 28 30 32 34 36 38 40 42 44 46 48 50

Fan Swirl Angle, deg

Blade Span, %

 Fan Speed

 61.7%

 75%

 87.5%

 100%

 Nozzle

 LF

 BL

 DP

 HF

Tip

Hub
NASA/TM—2013-214029 29

Figure 14.—Comparison of corrected thrust with fan nozzle area on 

fixed operating lines as measured with force balances.

Figure 15.—Change in corrected thrust with difference in fan nozzle area for several corrected fan speeds.

0

200

400

600

800

1000

1200

1400

1600

1800

2000

2200

2400

2600

30 40 50 60 70 80 90 100 110

Corrected Fan Speed, %

Corrected Thrust, lbf

Solid Sym -Total Thrust

Open Sym - Fan Thrust

 Nozzle

 BL

 DP

 HF

-12

-10

-8

-6

-4

-2

0

2

4

6

8

10

0 2 4 6 8 10 12

Change in Fan Nozzle Exit Area, %

Change in Corrected Thrust, %

Design Point nozzle

High Flow nozzle

Solid Sym -Total Thrust

Open Sym - Fan Thrust

 Fan Speed

 50%

 61.7%

 70%

 80%

 87.5%

 95%

 100%

Baseline nozzle

Increased

Thrust

Decreased

Thrust
NASA/TM—2013-214029 30

Figure 16.—Comparison in fan stage OAPWL with fan 

nozzle area as a function of corrected total thrust.

Figure 17.—Difference in OAPWL relative to the baseline 

fan nozzle as a function of corrected total thrust.

400 600 800 1000 1200 1400 1600 1800 2000 2200 2400 2600

Corrected Stage Thrust, lbf

115

120

125

130

135

140

145

Overall

Sound

Power

Level,

dB

Baseline nozzle

Design Point nozzle

High Flow nozzle

|

Nominal

Approach

|

Nominal

Cutback

|

Nominal

Takeoff

400 600 800 1000 1200 1400 1600 1800 2000 2200 2400 2600

Corrected Stage Thrust, lbf

-10

-5

0

5

Delta

Overall

Sound

Power

Level,

dB

Zero reference - Baseline nozzle

Design Point nozzle

High Flow nozzle

Increased

noise

Decreased

noise

nominal

approach

|

nominal

cutback

|

nominal

takeoff

|
NASA/TM—2013-214029 31

a) 61.7 percent fan design speed (designated approach speed)

b) 87.5 percent fan design speed (designated cutback speed)

Figure 18.—Sound power level spectra for acoustic 

rating speeds (5.9 Hz bandwidth) (continued).

0 1000 2000 3000 4000 5000 6000 7000 8000

Frequency, Hz

80

85

90

95

100

105

110

Sound

Power

Level,

dB

Baseline nozzle

Design Point nozzle

High Flow nozzle

BPF

|

2BPF

|

0 1000 2000 3000 4000 5000 6000 7000 8000

Frequency, Hz

90

95

100

105

110

115

120

125

Sound

Power

Level,

dB

Baseline nozzle

High Flow nozzle

High Flow nozzle

BPF

|
NASA/TM—2013-214029 32

c) 100 percent fan design speed (designated takeoff speed)

Figure 18.—Sound power level spectra for acoustic 

rating speeds (5.9 Hz bandwidth) (concluded).

Figure 19.—Percent change in fan stage OAPWL as a function of percent change in fan nozzle area.

0 1000 2000 3000 4000 5000 6000 7000 8000

Frequency, Hz

95

100

105

110

115

120

125

130

Sound

Power

Level,

dB

Baseline nozzle

Design Point nozzle

High Flow nozzle

BPF

|

-4

-3

-2

-1

0

1

2

0 2 4 6 8 10 12

Change in Fan Nozzle Exit Area, %

Change in Overall Sound Power Level (OASPL), %

 50%

 61.7%

 70%

 80%

 87.5%

 95%

 100%

Design Point nozzle

High Flow nozzle

Baseline nozzle
NASA/TM—2013-214029 33

a) Axial mean velocity

b) Tangential mean velocity

c) Axial turbulent velocity

d) Tangential turbulent velocity

Figure 20.—LDV mean velocity and turbulence measurements at 61.7 percent

corrected fan speed (approach conditions) for three fan nozzle configurations.

 

 

 

 

Baseline

Nozzle

Design Point

Nozzle

High Flow 

Nozzle

Blade Wake

Blade Wake
NASA/TM—2013-214029 34

a) Axial mean velocity

b) Tangential mean velocity

c) Axial turbulent velocity

d) Tangential turbulent velocity

Figure 21.—LDV mean velocity and turbulence measurements at 87.5 percent

corrected fan speed (cutback condition) for three fan nozzle configurations.

 

 

 

 

Baseline

Nozzle

Design Point

Nozzle

High Flow 

Nozzle
NASA/TM—2013-214029 35

a) Axial mean velocity

b) Tangential mean velocity

c) Axial turbulent velocity

d) Tangential turbulent velocity

Figure 22.—LDV mean velocity and turbulence measurements at 100 percent

corrected fan speed (takeoff condition) for three fan nozzle configurations.

 

 

 

 

 

 

 

0% 5% 7.5%

Baseline

Nozzle

Design Point

Nozzle

High Flow 

Nozzle
REPORT DOCUMENTATION PAGE Form Approved 

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1. REPORT DATE (DD-MM-YYYY)

01-12-2013 

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4. TITLE AND SUBTITLE

The Effect of Bypass Nozzle Exit Area on Fan Aerodynamic Performance and Noise in a 

Model Turbofan Simulator 

5a. CONTRACT NUMBER

5b. GRANT NUMBER

5c. PROGRAM ELEMENT NUMBER

6. AUTHOR(S)

Hughes, Christopher, E.; Podboy, Gary, G.; Woodward, Richard, P.; Jeracki, Robert, J. 

5d. PROJECT NUMBER

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5f. WORK UNIT NUMBER

WBS-22-781-30-75 

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John H. Glenn Research Center at Lewis Field 

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E-15384 

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Subject Category: 02 

Available electronically at http://www.sti.nasa.gov 

This publication is available from the NASA Center for AeroSpace Information, 443-757-5802

13. SUPPLEMENTARY NOTES

14. ABSTRACT

The design of effective new technologies to reduce aircraft propulsion noise is dependent on identifying and understanding the noise sources 

and noise generation mechanisms in the modern turbofan engine, as well as determining their contribution to the overall aircraft noise 

signature. Therefore, as part of the NASA Quiet Aircraft Technology program, a comprehensive aeroacoustic wind tunnel test program was 

conducted called the Fan Broadband Source Diagnostic Test. Performed in the NASA Glenn 9- by 15-Foot Low Speed Wind Tunnel, a 1/5 

scale model was used to simulate the bypass section of a current generation high bypass turbofan engine. The test objectives were to: 

identify the noise sources within the model and determine their noise contribution; investigate the impact of component design technology 

on noise level and aerodynamic performance; and conduct detailed flow diagnostics to characterize the physics of the noise generation 

mechanisms. This report discusses the results obtained investigating the effect of the bypass nozzle exit area on fan farfield noise and 

aerodynamic performance. The results will show that there were significant changes in aerodynamic performance and farfield acoustics as 

the fan nozzle area is increased: weight flow through the fan model increased between 7 and 9 percent; the combined fan stage adiabatic 

efficiency increased between 2 and 3 percent; total net thrust produced by the fan model increased between 2 and 3.5 percent; and the 

overall sound power farfield noise level dropped nearly linearly between 1 and 3.5 dB.

15. SUBJECT TERMS

Rotors; Fan blades; Noise reduction; Broadband; Wakes; Thrust; Turbofans; Acoustic properties; Aerodynamic characteristics; Nozzle 

flow; Turbulence; Simulators; Guide vanes; Stator; Microphones; Wind tunnel tests; Engine noise; Scale models 

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