Douglas C. Mathews and Larry A. Bock

United Technologies Corporation, Pratt & Whitney, East Hartford, Connecticut

Gerald W. Bielak, R.P. Dougherty, John W. Premo, and Dan F. Scharpf 

Boeing Commercial Airplane Group, Seattle, Washington

Jia Yu

Goodrich Aerostructures, Chula Vista, California

Pratt & Whitney/Boeing Engine Validation 

of Noise Reduction Concepts

Final Report for NASA Contract NAS3-97144, Phase 1

NASA/CR—2014-218088

February 2014
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Douglas C. Mathews and Larry A. Bock

United Technologies Corporation, Pratt & Whitney, East Hartford, Connecticut

Gerald W. Bielak, R.P. Dougherty, John W. Premo, and Dan F. Scharpf 

Boeing Commercial Airplane Group, Seattle, Washington

Jia Yu

Goodrich Aerostructures, Chula Vista, California

Pratt & Whitney/Boeing Engine Validation 

of Noise Reduction Concepts

Final Report for NASA Contract NAS3-97144, Phase 1

NASA/CR—2014-218088

February 2014

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared under Contract NAS3–97144, Phase 1
Acknowledgments

Acknowledgements are given to the following persons who comprised the core of the multi-company/government agency team: 

NASA Glenn Research Center, Project Manager: Dennis Huff; Task Manager: Naseem Saiyed; Pratt & Whitney, East Hartford, 

CT, Acoustics: Douglas C. Mathews, Robert J. Manzi, and Larry A. Bock; Boeing Commercial Airplane Group, Seattle, WA, 

Acoustics: Gerald W. Bielak, R. P. Dougherty, John W. Premo, and Dan F. Scharpf; Goodrich Aerostructures, Chula Vista, CA, 

Acoustics: Jia Yu; and MS, Newington, CT, Program Administration and Report Preparation: Robert D. Thulin.

Available from

NASA Center for Aerospace Information

7115 Standard Drive

Hanover, MD 21076–1320

National Technical Information Service

5301 Shawnee Road

Alexandria, VA 22312

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

Trade names and trademarks are used in this report for identification 

only. Their usage does not constitute an official endorsement, 

either expressed or implied, by the National Aeronautics and 

Space Administration.

Level of Review: This material has been technically reviewed by expert reviewer(s).
NASA/CR—2014-218088 iii

Forward

This report presents results of the work completed in Phase 1 of the Engine Validation of Noise 

Reduction Concepts (EVNRC) contract. The purpose of the program was to validate, through engine 

testing, advanced noise reduction concepts aimed at reducing engine noise up to 6 EPNdB and improving 

nacelle suppression by 50 percent relative to 1992 technology. Phase 2 of the program is currently near 

completion and upon its conclusion will be summarized in a separate report. 

Summary

The EVNRC program was authorized by NASA in August 1997 to validate, through engine testing, 

noise reduction concepts and technologies that have evolved during the AST program. A team comprised 

of members from NASA, Pratt & Whitney, Boeing and Goodrich Aerostructures (formerly Rohr) 

participated in bringing together the new hardware, software, test equipment and support personnel 

necessary to complete Phase I of these tests on a PW 4098 engine, the latest and largest of P&W’s 4000 

series engines. Testing was conducted at P&W’s Florida Test Facility on outdoor stand C-11 which is 

configured and instrumented to FAA noise measurement and data quality standards.

The major noise reduction concepts tested in this phase consisted of:

1. A new Boeing inlet featuring a scarfed design where the lower lip (keel) extends forward of the 

upper lip (crown) so that noise directivity is changed. This inlet also incorporates advanced triple 

layer liner treatment covering the entire internal surface from the mating flange of the engine’s 

front fan case forward to the inlet highlight without any hardwall splices and patches.

2. Advanced noise liner treatment developed by Goodrich for the engine’s front fan case. The 

treatment consisted of a 2-in. deep DDOF hybrid linear liner with a removable laser drilled 

polyurethane film bonded to the perforated face sheet. Removing the polyurethane allowed 

testing of a lower acoustical resistance liner than with the polyurethane installed.

3. By applying tape over various areas of the noise treatment, hardwall engine surfaces were 

simulated for testing purposes.

The major results of the testing were:

1. The new Boeing inlet with the advanced treatment in the front fan case showed a 2 to 4 EPNdB 

reduction in inlet fan noise at approach and cutback powers, compared to the production 

configuration. Approximately 80 percent dB nacelle suppression improvement was observed at 

approach power and 50 percent dB suppression improvement at cutback power.

2. A significant blade passing frequency (BPF) tone was measured with the scarf inlet, preventing 

an additional 1 EPNdB in noise reduction.

3. The BPF tone was found to be associated with the Boeing scarf inlet and was due to a flow 

separation problem at the crown location. This flow separation caused vortices to be shed into the 

fan. CFD calculations indicate that this separated flow will be eliminated with forward motion.

4. Noise treatment splices (joints) near the fan can produce large noise penalties of up to 4 EPNdB 

depending on the number and width of the splices.

5. Boeing’s RDIFF predictions for tone and broadband lining attenuation generally were within 

15 percent of the measured data in the direct radiation field but only within about 50 percent in 

the diffraction dominated field (aft of 70°). The very high frequency (8 to 10 kHz) attenuation by 

the lining was not accurately predicted.

6. Boeing’s phased array of 198 Kulite microphones mounted in the Inflow Control Device (ICD), 

effectively showed that at approach power the inlet noise of the PW4098 has a strong 

contribution from the low compressor. Fan blade passage tones and broadband fan noise are also 

important. The inlet lip lining was very effective in attenuating buzzsaw noise at low spinning 

orders.

NASA/CR—2014-218088 v

Contents

Forward........................................................................................................................................................iii

Summary......................................................................................................................................................iii

1.0 Introduction...........................................................................................................................................1

1.1 Background..................................................................................................................................1

1.2 EVNRC Program Description .....................................................................................................1

1.2.1 Phase 1............................................................................................................................2

1.2.2 Phase 2............................................................................................................................3

2.0 Noise Reduction Concepts and Hardware ............................................................................................3

2.1 Boeing Inlet .................................................................................................................................3

2.1.1 Scarf Inlet Design Objectives.........................................................................................3

2.1.2 Internal Aerodynamics Design .......................................................................................4

2.1.3 Scarf Inlet Pressurization Test........................................................................................8

2.1.4 Contour Confirmation.....................................................................................................8

2.1.5 Acoustic Linings...........................................................................................................11

2.1.6 Impedance Confirmation ..............................................................................................14

2.1.7 Conclusions ..................................................................................................................15

2.2 P&W Forward Fan Case ............................................................................................................16

2.3 Goodrich Case Liner.................................................................................................................16

2.3.1 Design Criteria..............................................................................................................16

2.3.2 Liner Configurations.....................................................................................................17

2.3.3 Predicted Liner Impedance...........................................................................................17

3.0 Test Program.......................................................................................................................................24

3.1 Test Site Characteristics.............................................................................................................24

3.1.1 Overall Site...................................................................................................................24

3.1.2 Engine Support Structure and Test Engine...................................................................25

3.1.3 Acoustical Arena ..........................................................................................................25

3.1.4 Acoustic Barriers..........................................................................................................25

3.2 Engine and Noise Measurement ................................................................................................27

3.2.1 Data Acquisition and Reduction System......................................................................27

3.2.2 Acoustical Calibrations.................................................................................................29

3.2.3 Test Procedures, Engine and Noise Data Acquisition ..................................................29

3.2.4 Test Data Processing Procedures for Far Field Real Time Data ..................................33

4.0 Test Results.........................................................................................................................................33

4.1 Far Field Data ............................................................................................................................33

4.1.1 Production Inlet Versus Hardwall Production Inlet—Baseline....................................34

4.1.2 Hardwalled Scarf Inlet Versus Hardwalled Production Inlet .......................................35

4.1.3 Benefits of Treated Scarf Inlet Plus FFCC...................................................................35

4.1.4 Forward Fan Case Treatment Benefits.........................................................................37

4.1.5 Scarf Inlet Diffuser Treatment Benefits........................................................................38

4.1.6 Inlet Lip Treatment Benefits.........................................................................................40

4.1.7 Fan Case Liner Splice Penalties ...................................................................................41

4.1.8 Inlet Liner Splice Penalties...........................................................................................44

4.1.9 Inlet Orientation Effects ...............................................................................................45

4.1.10 Scarf Effect on Aft Noise .............................................................................................45

4.1.11 Summary of Various Treatment Benefits.....................................................................47

4.1.12 Conclusions—Far Field Data .......................................................................................48

4.2 RDIFF Prediction Comparisons.................................................................................................49

4.2.1 RDIFF Prediction Code ................................................................................................49
NASA/CR—2014-218088 vi

4.2.2 Approach Power ...........................................................................................................49

4.2.3 Cutback Power..............................................................................................................54

4.2.4 Scarf Lining Variations ................................................................................................58

4.2.5 Conclusions ..................................................................................................................59

4.3 ICD Microphone Array Test Results.........................................................................................60

4.3.1 Background...................................................................................................................60

4.3.2 Instrumentation.............................................................................................................60

4.3.3 Phase Calibration..........................................................................................................66

4.3.4 Processing Methods......................................................................................................67

4.3.5 Test Configurations ......................................................................................................68

4.3.6 Results ..........................................................................................................................70

4.3.7 Conclusions ..................................................................................................................86

4.4 PW4098 Blade-Mounted Transducer and Mode Array Tone Diagnostic Test Results.............86

4.4.1 Data Acquisition and Instrumentation..........................................................................86

4.4.2 BMT Data Analysis......................................................................................................91

4.4.3 Flow Visualization......................................................................................................101

4.4.4 Inlet Mode Array Analysis.........................................................................................102

4.4.5 Conclusions and Recommendations...........................................................................112

5.0 Conclusions.......................................................................................................................................113

5.1 Far Field Data ..........................................................................................................................113

5.2 RDIFF Prediction Comparisons...............................................................................................114

5.3 ICD Microphone Array............................................................................................................114

5.4 BMT and Mode Array Tone Diagnostics ................................................................................115

Appendix...................................................................................................................................................117

References.................................................................................................................................................139
NASA/CR—2014-218088 1

Pratt & Whitney/Boeing Engine Validation 

of Noise Reduction Concepts

Final Report for NASA Contract NAS3-97144, Phase 1

Douglas C. Mathews and Larry A. Bock

United Technologies Corporation

Pratt & Whitney

East Hartford, Connecticut 06108 

Gerald W. Bielak, R.P. Dougherty, John W. Premo, and Dan F. Scharpf 

Boeing Commercial Airplane Group

Seattle, Washington 98124

Jia Yu

Goodrich Aerostructures

Chula Vista, California 91910

1.0 Introduction

1.1 Background

Major airports in the world’s air transportation systems face a serious problem in providing greater 

capacity to meet the ever increasing demands of air travel. This problem could be relieved if airports are 

allowed to increase their operating time, now restricted by curfews and by relaxing present limits on 

takeoffs and landings. The key operational issue in extending the present curfews is noise.

In response to these increasing restrictive noise regulations, NASA has launched a program to 

validate through engine testing, noise reduction concepts and technologies that have evolved from the 

Advanced Subsonic Technologies (AST) Noise Reduction Program. The goal of this AST program was to 

develop and validate technology that reduces engine noise and improves nacelle suppression effectiveness

relative to 1992 technology. Contract NAS3-97144 titled “Engine Validation of Noise Reduction 

Concepts” (EVNRC) was awarded to P&W on August 12, 1997 to conduct full scale noise reduction tests 

in two Phases on a PW4098 engine. The following Section 1.2 provides a brief description of the overall 

program. The remainder of this report provides a detailed documentation of Phase I of the program. 

1.2 EVNRC Program Description

The EVNRC program is a highly teamed effort involving NASA, Pratt & Whitney (P&W), Boeing, 

and Goodrich (formerly Rohr). The program duration was from August 1997 through September 2001 

and consisted of two phases. The two phases correspond to separate noise test programs in late 1998, and 

mid-2001 and were conducted with P&W’s latest and largest turbofan engine, the PW4098. This engine, 

which powers growth versions of Boeing’s 777 aircraft, has a 112.9-in. diameter fan and generates a rated 

thrust of 98,000 lb. The tests were planned to validate in full-scale a number of advanced noise reduction 

concepts and to demonstrate compliance with the engine noise reduction goals and nacelle aeroacoustic 

goals of the Advanced Subsonic Technology Program.

The program encompasses a very wide range of noise reduction technology such as:

• Advanced noise reduction treatment in a novel Boeing inlet 

• Advanced noise reduction treatment in P&W’s engine cases

• Low noise, cut on fan exit guide vanes

• Primary jet nozzle with acoustic lining to attenuate turbine noise
NASA/CR—2014-218088 2

The potential noise reduction from the technologies tested in this program was estimated to be in the 

2 to 4 EPNdB range at each condition, relative to the baseline PW4098 engine. Relative to 1992 

technology, noise reductions of up to 6 EPNdB in engine noise were expected, and nacelle suppression 

was expected to be improved by at least 50 percent.

This program offers high potential for technology transfer to commercial product lines. Reduced 

noise will be an internationally competitive issue, and this program is focused on technologies with a high 

probability of cost effective transition into the marketplace for new applications, as well as for retrofit or 

incorporation into today’s production engines. Specific and direct beneficiaries of the technology 

demonstrated in this proposed program will be the GP7000, the new engine being developed jointly by a 

P&W/General Electric (GE) Engine Alliance for installation into Boeing’s nacelle for the new Boeing 

747MD (Major Derivative) stretched aircraft. The program output will also be applicable to the remainder 

of Boeing’s nacelle product line, to the entire PW4000 engine family, to P&W’s new PW6000 mid-thrust 

engine, and to P&W’s future Advance Ducted Propulsor (ADP) family of engines, as well as other 

manufacturer product lines.

The program cost has been reduced because of the hardware sharing of team members. Boeing 

supplied the novel low noise inlet, valued at $1.5 million for the program, at no cost. Pratt & Whitney 

supplied the PW4098 engine and aft nacelle. Noise tests were conducted in conjunction with other 

planned P&W engine tests to reduce cost. Use of the P&W world-class outdoor C-11 test facilities in 

Florida, configured and instrumented to FAA noise measurement and data quality standards, also reduced 

costs.

Phase I has been completed and the purpose of this report is to document the content and results of 

that Phase. Phase 2 tests have been completed and data analysis is near completion. Results from Phase 2 

will be reported separately. 

1.2.1 Phase 1

The work of Phase I was broken into seven subtasks, described as follows:

1.2.1.1 Subtask 1.1.1—Concept Selection and Acoustic Design

The noise concepts that were defined by P&W and Boeing in the contract proposal were finalized in 

this subtask. In addition Goodrich was chosen as the supplier for the sound treatment liner material. The 

noise concepts selected were:

Boeing’s Ideal Inlet:

• An extended lower lip (scarfing) for noise shielding and noise redirection.

• Advanced liner design and construction that features triple layer broadband acoustic lining with 

minimum splice widths.

• Treatment that extends upstream beyond the inlet throat and highlight.

• Circumferentially uniform liner close to the fan.

• No probes and associated hardwalled patches.

P&W’s Engine Fan Case Treatment:

• A modified forward fan containment case that allowed testing of the current production liner, a 

hardwalled liner, and the new advanced linear double layer liner supplied by Goodrich, with only 

two circumferential segments separated by narrow hard wall splices.

• A design to minimize both community and interior buzzsaw noises.

These concepts are described more completely in Section 2.0 with diagrams and pictures.
NASA/CR—2014-218088 3

1.2.1.2 Subtask 1.1.2—Mechanical Design

This subtask provided for the development of working drawings suitable for the fabrication of the 

modified engine fan case hardware. Detailed design work for the Boeing inlet was completed prior to the 

contract.

1.2.1.3 Subtask 1.1.3—Fabrication

The fabrication subtask funded construction of the modified engine fan case and two spoolpieces, one 

spoolpiece incorporating the production Bill of Material (BOM) liner material and the other the advanced

Goodrich liner material.

1.2.1.4 Subtask 1.1.4—Boeing Subcontract

This subtask provided for Boeing’s participation to provide overall engineering support and 

consultation for the Phase I program including concept selection, test configuration selection, 

instrumentation system installation and operation during testing, data analysis, test results and reporting.

1.2.1.5 Subtask 1.1.5—Testing

The testing subtask covered costs associated with full scale PW4098 engine testing at P&W’s C-11 

Stand Acoustic Test Facility at West Palm Beach, Florida. Test costs were reduced by combining (piggybacking) the NASA funded tests with P&W funded PW4098 FAA Noise Certification Tests.

1.2.1.6 Subtask 1.1.6—Data Analysis

This subtask provided for the analysis of data collected during the testing subtask.

1.2.1.7 Subtask 1.1.7—Documentation

The parts of program documentation including preliminary test plans, design reviews with NASA, 

final test plans, design reports, data reviews with NASA, and final reports are covered by this subtask. 

1.2.2 Phase 2

The work of Phase 2 was also broken into the same seven subtasks as described for Phase 1. The

concepts chosen for testing in Phase 2 consisted of the Phase 1 concepts plus the following:

• Boeing’s scarf inlet fitted with a bellmouth lip instead of a flight lip.

• A cut on, 28 vane fan exit guide vane system.

• Fan blade number increased from 22 to 24 blades.

• Acoustically treated primary jet exhaust nozzle and plug.

2.0 Noise Reduction Concepts and Hardware

2.1 Boeing Inlet

2.1.1 Scarf Inlet Design Objectives 

Figure 1 summarizes the design objectives for the PW4098 demonstrator scarf inlet. The results of an 

earlier model scale scarf inlet test conducted at Boeing led to the decision to target a 15° scarf angle. It 

was desired to try to maximize the noise benefit of the scarf inlet by maximizing the scarf angle and at the 

same time push the comfort zone of the aerodynamics community while retaining a reasonable probability 

of meeting their requirements when tested. It was thought that a 15° scarf would meet this objective.

Propulsion aerodynamics experts were concerned about inlet recovery, fan face flow distortion and high 

angle of attack separation. External aerodynamics for cruise performance was also an issue because the 

propulsion concerns affected the lip design on the external cowl as well as internally. For the design 
NASA/CR—2014-218088 4

Figure 1.—Demonstrator scarf design objectives.

study, the engine outer cowl and internal fan case geometry were fixed so that the scarf geometry would 

have to fare into these lines in an aerodynamically satisfactory manner.

It was desired to use a flight lip on the inlet for the static test rather than a bellmouth lip that is usually 

used for static engine noise testing. The purpose of the bellmouth lip is to match the internal inlet flow more 

closely to flight by reducing the high lip wall flow Mach number that occurs with a flight lip in static test. 

The lip design is a major aerodynamics issue with the scarf inlet and is strongly interdependent on the inlet 

length and scarf angle. Boeing wanted to insure that these interdependencies were considered for this design 

as they would be for a product design. Also, it is known that the lip shape affects the sound radiation 

directivity. Since the purpose of the scarf shape is to change the radiation directivity a bellmouth lip could 

confound the interpretation of the test results. Unfortunately the wall Mach number increase associated with 

the flight lip in the static condition may also affect the radiation. It was felt that this would be a secondary 

affect at the approach engine power condition but there was concern about the high power conditions. 

A major acoustic objective was to acoustically treat the entire internal surface of the inlet and fan case 

from the fan rub strip to the inlet highlight without any hardwall splices and patches. Today’s production 

inlets have no acoustic treatment on 30 to 50 percent of the internal inlet surface because of de-icing, 

concerns about boundary layer growth near the inlet throat, panel seams and structural reinforcement. 

Fully treating all surfaces of the scarf inlet would allow these hardwall surfaces to be simulated by taping 

over the lining to measure their acoustic importance. In addition it was decided to use advanced concept 

triple layer acoustic linings within the inlet. It was realized that the surface curvatures within the inlet 

would make building a triple layer lining with minimal splices a challenge but Boeing’s Wichita 

manufacturing personnel believed it could be done for a static test inlet. 

2.1.2 Internal Aerodynamics Design

2.1.2.1 Lines

Scarf inlet contours along the crown and keel are compared with the PW4098 production inlet 

contours in Figure 2. Two scarf inlet designs are represented. One is for the 15° objective and the second 

is a biplanar design with a 15° scarf on the bottom of the inlet and a smaller 12° scarf on the top, for a net 

13° scarf angle. The biplanar scarf design was necessitated by the results of a Navier Stokes computer 

internal flow simulation of the 15° design. The Navier Stokes (NS) calculations were run for final 

confirmation of the inlet design iterations which were done using potential flow codes for max engine 

power inlet mass flows. 
NASA/CR—2014-218088 5

Figure 2.—Comparison of production and scarf inlet contours.

2.1.2.2 Distortion

The Navier Stokes (NS) results for the three inlets represented in Figure 2 are shown in Figure 3. For 

max power static operation, the Navier Stokes calculation predicted that the 15° scarf inlet would result in 

flow separation near the inlet crown and an associated distortion at the fan face. While it is expected that 

the separation would clean up with forward motion, it was determined that it may reduce the useful life of 

the fan blades given the amount of time the engines are run at high power with the airplane at rest. In 

addition, a calculation of the acoustic effect of the distortion concluded that it would generate a high fan 

tone noise level. The biplanar concept was therefore invented as a simple way to maintain much of the 

tooling that had been built by the time the NS results were completed and to solve the distortion problem. 

Figure 3 shows that the NS calculation for the biplanar scarf predicted a much smaller distortion at 

the fan face at max power. The minimal distortion predicted at high engine power for the biplanar scarf 

led to the expectation that the distortion would totally disappear at approach engine power and perhaps 

cutback engine power as well so that distortion related fan tone generation would not be a serious 

problem. NS simulations for these engine powers were not done because of the high cost of these 

calculations.
NASA/CR—2014-218088 6

Figure 3.—Navier Stokes simulation—fan face distortion, PW4098, max power.

2.1.2.3 Mach Number Contours and Boundary Layer Thickness

Wall Mach number contours were calculated at approach, cutback and takeoff engine powers with a 

3D potential flow code for the 13° biplanar scarf. The lower power calculations were not part of the 

design process and were not done until after manufacturing of the inlet had begun. The calculations were 

done for the crown (0°), mid line (90°) and keel (180°) of the inlet. For comparison the contours were 

calculated for both static operation (noise test situation) and flight. The results of these calculations are 

shown in Figure 4. The most striking feature of these results is the difference between static and flight.

The “sink” flow turning around the inlet lip at the static condition results in very high Mach numbers near 

the inlet lip. The entire lip flow is supersonic at takeoff and the crown is supersonic at cutback. In flight 

this turning is reduced significantly and the Mach numbers are much lower. The supersonic flow in the 

inlet was of concern since it will have an effect on the propagation of noise and not be representative of 

flight. It was felt that the higher subsonic speed flow would not be a significant problem at the approach 

power condition but cutback and takeoff were of concern.

Figure 5 shows estimates of the fan face boundary layer thickness at the same azimuth angles for 

which the axial Mach number contours shown above were calculated. These calculations were done with 

a 3D boundary layer code. It is seen that the static boundary layer thickness at the crown, in particular, is 

much larger than for the flight condition. The thickness increase also increases with engine power. The 

azimuth spatial definition of the boundary layer thickening at the crown was not determined.
NASA/CR—2014-218088 7

Figure 4.—Wall Mach number profiles for PW4098 demonstrator scarf inlet.

Figure 5.—Fan face boundary layer thickness, scarf inlet, static versus flight.
NASA/CR—2014-218088 8

2.1.3 Scarf Inlet Pressurization Test

The inlet pressurization test was used to evaluate the potential sources of BPF tone due to leakage of 

air through the inlet liner. Covering the access holes with aluminum plates and applying RTV on all of the 

seams sealed the rear bulkhead of the inlet. The inlet was filled with shop air to a pressure of 2 psig and 

inspected for leaks. Small leaks were found in the bulkhead, and in the inlet where the inlet-probe plug 

was installed. Air passage through these leaks was barely perceptible by touch, and hence was not 

considered to be a contributor to the generation of the BPF tone.

2.1.4 Contour Confirmation

2.1.4.1 Lip Contour

Measurements were made to verify the lip contour of the inlet using templates placed at the crown, 

±90° from the crown and at the keel. Measurements were made from the inlet highlight back 27 in. into 

the inlet every 0.5 in. Nearly 90 percent of the measurements were within 40 mil of the specified contour 

with a maximum deviation of 85 mil. This was considered to be within necessary tolerance.

2.1.4.2 “A” Flange Contour

An inlet template (Figure 6) was used to verify the circular uniformity of the inlet at the attach flange. 

The template was used with a “banana gauge” in order to quantify the deviations. Since the template only 

covered 90° of the inlet circumference, it was moved around the inlet at 45° intervals, while the gap 

between the template and inlet wall was measured every 5° relative to the template center. The template 

was designed with “feet” at each end in order to provide a gap between the template and the inlet. This 

also created a different reference for each measurement location. Because of this, no absolute 

measurement of the circumferential variation was possible by this technique.

Figure 7 shows the measurements from the template for the scarf inlet. Measurements were taken 

with the inlet de-pressurized and on the ground. By following a particular line style (solid or dashed) the 

variation can be seen as though the template were placed end-to end with no overlap. Differences between 

the solid and dashed curves provide some indication as to the variations in gap due to the different 

reference heights for each template location. The average measurement at each angular position is shown 

in Figure 8. A regular or periodic variation in inlet radius was not evident from these measurements. 

Figure 6.—Inlet template used to measure circumferential 

uniformity at the attach flange.
NASA/CR—2014-218088 9

Figure 7.—Measured variations of the gap between the template and the scarf inlet at the attach flange.

Figure 8.—Averaged variations of the gap between the template and the scarf inlet at the 

attach flange.

A Fourier decomposition of the average, template-measured displacements is shown in Figure 9. The 

mode numbers represent physical distortions of the inlet with a circumferential period equal to the mode 

number; e.g., an n = 4 mode implies that a four-lobe distortion of the circumference exists around the 

inlet. This type of physical distortion would create local steady distortions in the inlet boundary layer that 

could interact with the fan blades causing a BPF tone. The effect of this physical distortion may also be 

amplified when the inlet is attached to the fan case if the fan case is also not perfectly uniform. Figure 9

suggests that there are no dominant distortion modes. 

Since relative measurements were used instead of absolute measurements, a numerical study was 

conducted to determine the error associated with using relative measurements to calculate distortion 

orders. Figure 10 shows the mode amplitudes for a 100-point pseudo-random distortion, which was 

analyzed in a manner similar to the physical measurements. Figure 10 shows the resulting mean and 

95 percent confidence error for this study. The conclusion is that using relative measurements allows an 

accurate assessment of the higher distortion orders, but large errors exist at the lower distortion orders.
NASA/CR—2014-218088 10

Figure 9.—Fourier decomposition of the averaged variations of the gap between the 

template and the scarf inlet at the attach flange.

Figure 10.—Numerical experiment showing error associated with using relative 

measurements.

When the inlet was mounted on the production fan case on the engine, measurements were made of 

the fit of the inlet to the case. The size of the internal wall step from the inlet to the fan case was measured 

at 5° intervals around the entire inlet. The results of these measurements are shown in Figure 11(a), 

positive indicating a downward step from the inlet to the fan case. A step of approximately 0.1 in was 

found on the top part of the inlet. This step was smoothed to some degree at the test site with silicon 

adhesive. The Fourier components of the contour variations are also shown in Figure 11(b). A flow 

distortion associated with this shape variation would not be a significant cause of fan tone noise since the 

maximum modal amplitudes are in the lowest order components. Also the 0.1 in step is small compared to 

the 2 to 3 in. boundary layer thickness in the crown region so it is probably insignificant for broadband 

noise generation as well.
NASA/CR—2014-218088 11

Figure 11.—(a) Displacement of inlet at “A” flange relative to engine fan 

case, and (b) Scarf “A” flange displacement spectrum

2.1.5 Acoustic Linings

Three different acoustic lining designs were built for the scarf inlet and two designs were used for the 

advanced forward fan spool. Figure 12 shows schematics of these designs and where in the inlet they 

were used. Triple layer liners were used for the inlet and double layer liners used for the fan case spool. 

(Liner information for the P&W/Goodrich fan case is discussed in more detail in Section 2.3.) Liners with 

smaller open area face sheets were used in the regions of the inlet with higher Mach numbers. This was 

done to reduce “self-noise” generated as the high-speed air passed over the lining orifices. The diffuser 

lining was designed to an impedance of (2.6, –0.5) at the approach condition. Figure 13 shows the 

predicted acoustic impedances of the inlet liners at the approach and cutback conditions. The high grazing 

flow Mach numbers at the lip, together with the smaller open area of the lip liner face sheet, results in the 

relatively high acoustic resistances predicted for the lip liner and, to a lesser extent, for the forward 
NASA/CR—2014-218088 12

diffuser liner. The impedances at cutback engine power were a fall out of the designs of the liners that 

were optimized for the approach condition. The design target impedances for the approach and cutback 

conditions are shown for reference. A study was done optimizing the liners for the cutback condition with 

the approach power impedance then being a fall out. These liners were significantly inferior to the liners 

designed for approach when both conditions are considered.

Figure 14 shows the predicted impedances of the advanced fan case spool liners which are discussed 

in more detail in Section 2.3. Two sets of impedances are shown for each engine power condition. The 

higher resistances were obtained by placing a layer of 7.2 percent open area perforated polyurethane tape 

over the 32 percent open area punched aluminum face sheet of the liner. The tape is approximately 15 mil 

thick with 5 mil diameter holes. The spool piece liner was designed to have only two hardwall 0.1 in. 

splices, one at the top and one at the bottom of the spool. In actuality the splices were about 0.375 in. 

wide.

Figure 12.—Acoustic lining details.
NASA/CR—2014-218088 13

Figure 13.—Predicted impedance spectra of specified scarf inlet linings.
NASA/CR—2014-218088 14

Figure 14.—Impedance of advanced fan case liners.

2.1.6 Impedance Confirmation

Impedance measurements were made on the completed scarf inlet to verify that the acoustic linings 

behaved as designed (Figure 15). Measurements were made at approximately 200 locations distributed 

over the inlet with about 100 measurements within the aft diffuser. The measurements were made with a 

3 cm tube “portable impedance meter” similar to that conventionally used for lining samples in the 

laboratory. The primary difference between the portable device and the laboratory device is that the 

portable device uses a rubber seal to define the lining measurement face sheet area and the laboratory 

device cuts the lining test sample out of the panel and contains it within the impedance tube (Figure 16).

As shown in Figure 16, the cavity volumes outside the edge defined by the seal influence the 

measurement by the portable device. The measured data therefore must be compared to a calculation of 

the impedance of the lining segment that accounts for this edge effect. Considered in this way the test data 

showed impedance variations of ±0.5 rhoc for a given lining design over the inlet and the average 

impedance differed from that of the specified lining by ±0.5 rhoc depending on frequency. It was 

estimated that the average impedance deviation from the specified lining would have no effect on the inlet 

component EPNL and the variation could only have a worst case variation of ±0.5 EPNdB.
NASA/CR—2014-218088 15

Figure 15.—Acoustic impedance verification.

Figure 16.—Portable impedance meter.

2.1.7 Conclusions

The decision to use a flight lip resulted in high Mach number flow near the lip and thicker boundary 

layer at the fan face relative to flight.

The expected flow deviations from flight get smaller as engine power is reduced. It was therefore 

expected that they would not be acoustically significant at approach power. However, cutback power was 

a concern. The inlet lip contours were within specification. The inlet-fan case interface was out of 

specification but was probably not significant for fan tone generation or for broadband noise.

The measured variation of acoustic impedance was ±1⁄2 rhoc with a 0 to 1 rhoc shift in the mean 

depending on frequency. This was not expected to be significant.
NASA/CR—2014-218088 16

Figure 17.—PW4098 forward fan containment case—modified to incorporate new treatment.

2.2 P&W Forward Fan Case

The new forward fan case provided for this program has the capability to accept alternate liners 

forward of the fan. The case was modified to accept bolt-on spool pieces that featured the increased 

acoustic area and thickness associated with the advanced liners, as well as the production PW4098 

acoustic treatment. In order to accommodate these spool pieces, the fan case forward flange was relocated 

aft by the width of the spool piece and the Kevlar fan blade containment material was removed. The spool 

pieces were designed with a larger diameter inlet bolt circle than the production design at the A-flange 

engine/inlet interface in order to house the increased thickness treatment. A schematic view of the lower 

half of the modified fan case is shown in Figure 17. Two spoolpieces were designed and built by P&W, 

one which incorporated the P&W supplied PW4098 production treatment and the other featuring the 

advanced Goodrich treatment.

2.3 Goodrich Case Liner

2.3.1 Design Criteria

Two acoustic liner design targets were specified for the PW4098 forward fan case (Figure 17). The 

first design target was to match Boeing's scarf inlet diffuser's optimum liner impedance criterion. In this 

criterion, the average optimum impedance at the approach condition is Z/ρc = (R + jX) = 2.6 – j0.5 and 

for the cutback condition is Z/ρc = 3.3 – j0.8 (See Table 1 for the definitions of symbols). The range of 

design frequencies is from 630 to 3150 Hz to cover both BPF (blade passage frequency) tone, BPF 

derivatives and broadband noise. Since optimum impedance is a function of in-duct Mach number (higher 

Mach number has higher resistance component), the diffuser's optimum liner impedance near the A1 

flange is about Z/ρc = 2.0 – j0.5 for approach and Z/ρc = 3.1 – j0.8 for cutback. The Mach number in the 

fan case section is very close to the inlet diffuser section near A1 flange. Therefore, the optimum 

impedance with a lower resistance component was chosen as the first forward fan case liner design target.

The second design target is for cutback BPF and MPT noise suppression. The optimum impedance is 

Z/ρc = (2.0 – j1.0) at 1000 Hz. In addition, the acoustic liner must have uniform impedance properties and 

no more than two axial splices. The splice width was to be no more than 1/2 in.
NASA/CR—2014-218088 17

TABLE 1.—DEFINITIONS OF SYMBOLS

Z/ρc a complex number representing normalized impedance 

R/ρc the normalized acoustic resistance

j is √-1 (imaginary number)

X/ρc the normalized acoustic reactance

ρ the air density and c is the speed of sound 

ρc the characteristic impedance (unit: cgs - rayl)

2.3.1.1 Design Input Parameters 

Boeing and Pratt & Whitney (PWA) provided design parameters, which include in-duct spectra, 

Mach number, and boundary layer thickness at approach and cutback conditions. Since the design 

parameters were based on estimated data, the spectra provided by Boeing and PWA were slightly 

different. Figure 18 and Figure 19 show the details of all input parameters.

2.3.2 Liner Configurations

Rohr, Inc., developed an innovative design concept to achieve both design targets with minimum 

additional cost for the fan case test panel fabrication. The design is a DDOF hybrid linear liner, which is 

composed of a removable laser drilled polyurethane (PU) film bonded on a perforate face sheet, a fine 

non-metallic mesh septum, solid back skin, and two 3/8 in. size aluminum core blankets bonded between 

perforate plate sheet, mesh septum, and solid back skin. This DDOF liner not only provides necessary 

structural strength but also ensures a near locally reactive acoustic behavior. The DDOF mesh septum 

liner with laser drilled PU film, which provides a higher acoustic resistance component, is designed to 

meet the first target. When the PU film is removed from the DDOF liner surface, the liner is designed to 

meet the second target, which has a lower acoustic resistance component. Table 2 lists the design 

configurations with and without PU-film. Figure 20 is a cross section drawing of the liner. It shows that 

the depth of upper core (between face sheet to septum) is a constant 0.45-in. However, the depth of lower 

core (between septum and back skin) varies slightly with the average value being 1.60 in. The minimum 

total core depth is 2 in.

TABLE 2.—ACOUSTIC LINER DESIGN CONFIGURATIONS FOR FORWARD FAN CASE 

Liner Configuration with PU film Liner Configuration without PU film

PU Surface:

Pu-POA = 18; diam. = 0.0062 in.; thick = 0.015 

in.

Perforate sheet:

Pf-POA = 32; diam. = 0.05 in.; thick = 0.032 in.

Upper Core depth: 0.45 in. 

Mesh Septum:

R(105) = 90 rayl and NLF = 1.3

Lower core depth = 1.60 in.

PU Surface:

PU film removed.

Perforate sheet:

Pf-POA = 32; diam. = 0.05 in.; thick = 0.032 in.

Upper Core depth: 0.45 in. 

Mesh Septum:

R(105) = 90 rayl and NLF = 1.3

Lower core depth = 1.60 in.

2.3.3 Predicted Liner Impedance

Figure 21 to Figure 24 show the predicted acoustic impedance at approach and cutback conditions 

with and without PU film. Since PWA and Boeing have slightly different estimate spectra, the predicted 

acoustic impedance levels also show a small difference. The liner configuration with laser drilled PU film 

is designed for the first target. 
NASA/CR—2014-218088 18

At the approach condition, Figure 21 indicates that resistance components derived from either Boeing or 

PWA spectrum are near the design target of 2.0ρc at a frequency range up to 3150 Hz. The reactance 

components derived from either Boeing or PWA spectrum are also close to the –0.5ρc target value at 

frequencies between 1250 to 3150 Hz. At the cutback condition, Figure 22 shows that the resistance is 

slightly higher than target value by using PWA’s spectrum, but it hit the design target if Boeing's spectrum 

is used. The reactance components are near the design target at the frequency range from 1000 to 3150 Hz.

The liner configuration without laser drilled PU film (film removed) is designed for the second target. 

Figure 24 indicates that both resistance and reactance components are near the design target (2.0 – j1.0ρc) 

at 1000 Hz frequency.

Approach-0.3M

FREQ. PWA-SPL BAC-SPL

50.0 121.1 115.5

63.0 121.8 123.5

80.0 122.3 122.5

100.0 119.2 120.0

125.0 119.4 120.0

160.0 121.4 121.0

200.0 122.4 123.0

250.0 121.8 122.5

315.0 122.8 122.5

400.0 123.6 123.0

500.0 124.2 123.5

630.0 132.6 133.5

800.0 128.8 126.5

1000.0 128.0 123.5

1250.0 133.0 127.5

1600.0 132.0 126.5

2000.0 134.6 128.5

2500.0 135.1 128.0

3150.0 135.0 130.0

4000.0 135.4 132.0

5000.0 136.0 133.0

6300.0 135.8 133.0

8000.0 136.1 133.0

10000.0 132.1 129.0

CM = 0.300 TMP = 527.0 R 

SPR = 13.7 in., BL = 0.175 in.

Figure 18.—Forward fan case design input parameters at approach condition.
NASA/CR—2014-218088 19

Cutback-0.48M

FREQ. PW-SPL BAC-SPL

50.0 135.3 122.0

63.0 136.0 137.0

80.0 137.1 138.0

100.0 137.9 140.0

125.0 139.2 140.0

160.0 136.0 137.0

200.0 141.8 137.5

250.0 147.1 139.5

315.0 149.3 145.5

400.0 149.9 145.0

500.0 150.0 145.0

630.0 152.5 145.0

800.0 152.5 146.5

1000.0 159.3 156.5

1250.0 149.6 144.5

1600.0 150.3 143.5

2000.0 148.3 144.0

2500.0 144.9 138.5

3150.0 141.2 137.0

4000.0 136.3 133.0

5000.0 133.0 132.0

6300.0 131.1 131.5

8000.0 130.7 131.0

10000.0 126.9 126.6

CM = 0.480 TMP = 513.0 R 

SPR = 12.5 in., BL = 0.175 in.

Figure 19.—Forward fan case design input parameters at cutback condition.

Figure 20.—Forward fan case liner configuration sketch.
NASA/CR—2014-218088 20

Approach-0.3M

FREQ. PWA-SPL RESIS REAC FREQ. BAC-SPL RESIS REAC

50.0 121.1 2.12 -20.99 50.0 115.5 1.99 -20.99

63.0 121.8 2.12 -16.66 63.0 123.5 1.99 -16.66

80.0 122.3 2.12 -13.11 80.0 122.5 1.99 -13.11

100.0 119.2 2.12 -10.49 100.0 120.0 1.99 -10.48

125.0 119.4 2.12 -8.38 125.0 120.0 1.99 -8.38

160.0 121.4 2.12 -6.54 160.0 121.0 1.99 -6.54

200.0 122.4 2.12 -5.22 200.0 123.0 1.99 -5.22

250.0 121.8 2.12 -4.16 250.0 122.5 1.99 -4.16

315.0 122.8 2.12 -3.29 315.0 122.5 1.99 -3.28

400.0 123.6 2.12 -2.56 400.0 123.0 2.00 -2.56

500.0 124.2 2.13 -2.02 500.0 123.5 2.00 -2.02

630.0 132.6 2.13 -1.57 BPF 630.0 133.5 2.01 -1.57

800.0 128.8 2.14 -1.20 800.0 126.5 2.02 -1.19

1000.0 128.0 2.16 -0.91 1000.0 123.5 2.03 -0.90

1250.0 133.0 2.18 -0.67 2BPF 1250.0 127.5 2.06 -0.66

1600.0 132.0 2.21 -0.47 1600.0 126.5 2.09 -0.45

2000.0 134.6 2.26 -0.35 3BPF 2000.0 128.5 2.15 -0.33

2500.0 135.1 2.30 -0.33 2500.0 128.0 2.20 -0.31

3150.0 135.0 2.20 -0.56 3150.0 130.0 2.11 -0.56

4000.0 135.4 1.02 -0.36 4000.0 132.0 0.91 -0.36

5000.0 136.0 1.29 0.79 5000.0 133.0 1.18 0.80

6300.0 135.8 1.46 1.02 6300.0 133.0 1.36 1.02

8000.0 136.1 1.09 1.54 8000.0 133.0 0.98 1.54

10000.0 132.1 1.56 2.40 10000.0 129.0 1.46 2.39

Figure 21.—Liner impedance at approach condition with PU film.
NASA/CR—2014-218088 21

Ct-0.48M

FREQ. PW-SPL RESIS REAC FREQ. BAC-SPL RESIS REAC

50.0 135.3 3.61 -20.73 50.0 122.0 3.11 -20.72

63.0 136.0 3.61 -16.45 63.0 137.0 3.11 -16.45

80.0 137.1 3.61 -12.96 80.0 138.0 3.11 -12.95

100.0 137.9 3.61 -10.37 100.0 140.0 3.11 -10.36

125.0 139.2 3.61 -8.30 125.0 140.0 3.11 -8.29

160.0 136.0 3.61 -6.49 160.0 137.0 3.11 -6.48

200.0 141.8 3.61 -5.20 200.0 137.5 3.11 -5.18

250.0 147.1 3.61 -4.17 250.0 139.5 3.11 -4.15

315.0 149.3 3.61 -3.32 315.0 145.5 3.11 -3.29

400.0 149.9 3.60 -2.62 400.0 145.0 3.11 -2.60

500.0 150.0 3.60 -2.11 500.0 145.0 3.11 -2.08

630.0 152.5 3.59 -1.69 630.0 145.0 3.10 -1.65

800.0 152.5 3.58 -1.36 800.0 146.5 3.10 -1.31

1000.0 159.3 3.56 -1.11 BPF 1000.0 156.5 3.09 -1.05

1250.0 149.6 3.54 -0.92 1250.0 144.5 3.08 -0.85

1600.0 150.3 3.49 -0.76 1600.0 143.5 3.05 -0.68

2000.0 148.3 3.42 -0.65 2BPF 2000.0 144.0 3.02 -0.58

2500.0 144.9 3.32 -0.58 2500.0 138.5 2.94 -0.53

3150.0 141.2 3.10 -0.59 3150.0 137.0 2.73 -0.59

4000.0 136.3 2.16 -0.26 4000.0 133.0 1.74 -0.26

5000.0 133.0 2.47 0.73 5000.0 132.0 2.06 0.75

6300.0 131.1 2.53 1.05 6300.0 131.5 2.13 1.04

8000.0 130.7 2.24 1.62 8000.0 131.0 1.82 1.62

10000.0 126.9 2.65 2.50 10000.0 126.6 2.25 2.48

Figure 22.—Liner impedance at cutback condition with PU film.
NASA/CR—2014-218088 22

App-0.3M

FREQ. PWA-SPL RESIS REAC FREQ. BAC-SPL RESIS REAC

50.0 121.1 1.45 -21.00 50.0 115.5 1.42 -21.00

63.0 121.8 1.45 -16.67 63.0 123.5 1.42 -16.67

80.0 122.3 1.45 -13.13 80.0 122.5 1.42 -13.13

100.0 119.2 1.45 -10.50 100.0 120.0 1.42 -10.50

125.0 119.4 1.45 -8.40 125.0 120.0 1.43 -8.40

160.0 121.4 1.45 -6.56 160.0 121.0 1.43 -6.56

200.0 122.4 1.46 -5.25 200.0 123.0 1.43 -5.25

250.0 121.8 1.46 -4.20 250.0 122.5 1.43 -4.20

315.0 122.8 1.46 -3.34 315.0 122.5 1.43 -3.33

400.0 123.6 1.46 -2.63 400.0 123.0 1.43 -2.62

500.0 124.2 1.46 -2.10 500.0 123.5 1.43 -2.10

630.0 132.6 1.47 -1.67 BPF 630.0 133.5 1.44 -1.67

800.0 128.8 1.48 -1.33 800.0 126.5 1.45 -1.32

1000.0 128.0 1.49 -1.07 1000.0 123.5 1.46 -1.06

1250.0 133.0 1.50 -0.88 2BPF 1250.0 127.5 1.48 -0.86

1600.0 132.0 1.53 -0.73 1600.0 126.5 1.51 -0.71

2000.0 134.6 1.56 -0.67 2000.0 128.5 1.55 -0.65

2500.0 135.1 1.58 -0.71 2500.0 128.0 1.59 -0.70

3150.0 135.0 1.45 -1.00 3150.0 130.0 1.47 -1.01

4000.0 135.4 0.30 -0.92 4000.0 132.0 0.30 -0.94

5000.0 136.0 0.55 0.09 5000.0 133.0 0.55 0.08

6300.0 135.8 0.69 0.15 6300.0 133.0 0.70 0.13

8000.0 136.1 0.32 0.45 8000.0 133.0 0.31 0.42

10000.0 132.1 0.74 1.06 10000.0 129.0 0.74 1.03

Figure 23.—Liner impedance at approach condition without PU film
NASA/CR—2014-218088 23

Cut-0.48M

FREQ. PW-SPL RESIS REAC FREQ. BAC-SPL RESIS REAC

50.0 135.3 2.01 -20.74 50.0 122.0 1.86 -20.73

63.0 136.0 2.01 -16.47 63.0 137.0 1.86 -16.46

80.0 137.1 2.01 -12.98 80.0 138.0 1.86 -12.97

100.0 137.9 2.01 -10.40 100.0 140.0 1.86 -10.39

125.0 139.2 2.01 -8.33 125.0 140.0 1.86 -8.32

160.0 136.0 2.01 -6.53 160.0 137.0 1.86 -6.51

200.0 141.8 2.01 -5.25 200.0 137.5 1.86 -5.23

250.0 147.1 2.01 -4.23 250.0 139.5 1.86 -4.21

315.0 149.3 2.00 -3.40 315.0 145.5 1.86 -3.37

400.0 149.9 2.00 -2.73 400.0 145.0 1.85 -2.69

500.0 150.0 1.99 -2.25 500.0 145.0 1.85 -2.19

630.0 152.5 1.97 -1.86 630.0 145.0 1.84 -1.80

800.0 152.5 1.94 -1.57 800.0 146.5 1.83 -1.48

1000.0 159.3 1.90 -1.37 BPF 1000.0 156.5 1.81 -1.27

1250.0 149.6 1.84 -1.23 1250.0 144.5 1.77 -1.12

1600.0 150.3 1.75 -1.12 1600.0 143.5 1.72 -1.01

2000.0 148.3 1.63 -1.06 2BPF 2000.0 144.0 1.65 -0.96

2500.0 144.9 1.48 -1.03 2500.0 138.5 1.54 -0.96

3150.0 141.2 1.23 -1.04 3150.0 137.0 1.31 -1.05

4000.0 136.3 0.39 -0.82 4000.0 133.0 0.38 -0.84

5000.0 133.0 0.66 -0.01 5000.0 132.0 0.66 0.00

6300.0 131.1 0.69 0.17 6300.0 131.5 0.70 0.14

8000.0 130.7 0.42 0.52 8000.0 131.0 0.40 0.49

10000.0 126.9 0.75 1.15 10000.0 126.6 0.77 1.10

Figure 24.—Liner impedance at cutback condition without PU film.

2.3.3.1 DC Flow Resistance Data

The DC flow resistance measurements for the liner septum were conducted at airflow rates of 30, 60, 

105, 150, and 200 cm/sec. All the data were normalized to reference ambient conditions (70 °F and 29.92 in. 

Hg). The first order least squares curve fit was used to generate required data R(105), and NLF. The R(105) is 

DC flow resistance data at 105 cm/sec and the NLF, which is referred as non-linear factor, is the ratio of 

resistance data at 200 cm/sec to data at 20 cm/sec (R(200)/R(20)). Table 3 shows the measured DC flow 

resistance of the actual test article. The data indicates that the acoustic test panel is within 90 ± 20 rayl 

specification.
NASA/CR—2014-218088 24

TABLE 3.—MEASURED SEPTUM DC FLOW RESISTANCE

Section 1 Section 2 Section 3 Section 4

1 109.7 118.5 88.5 97.2

2 117.6 121.2 90.9 92.1

3 106.1 100.4 88.3 95.6

4 111.2 104.6 81.0 91.2

5 101.6 96.4 91.2 87.5

6 107.3 103.5 84.1 82.0

7 106.0 99.2 89.1 82.3

8 106.0 98.2 88.8 83.2

9 99.6 97.9 93.0 81.4

10 103.5 98.4 94.5 83.0

11 104.8 99.0 94.3 88.0

12 113.7 105.0 85.5 87.8

13 102.6 105.0 90.5 80.7

14 110.2 99.4 84.3 84.7

15 98.0 101.0 98.3 87.3

16 97.8 101.3 82.8 81.9

17 95.6 97.4 94.1 87.5

18 104.7 109.1 87.8 81.9

19 101.0 97.1 93.2 84.0

20 100.8 100.4 82.6 87.2

21 96.0 98.2 96.2 82.3

22 98.0 102.9 82.9 87.8

23 113.4 106.3 95.6 94.6

24 110.9 123.1 94.0 92.2

Average

R(105) 104.8 103.5 89.63 86.8

NLF 1.32 1.32 1.37 1.28

Specification: 

Average Septum DC flow Resistance measured at 105 cm/sec. R(105) = 90 ± 20 RAYLS

Individual Septum DC flow Resistance measured at 105 cm/sec. R(105) = 90 ± 35 RAYLS 

NLF = R(200)/R(20) < 1.5 

2.3.3.2 PU Film and Perforate POA

The measurements of POA and hole diameter on the final test panels were not performed. However, 

the small coupon sample measurement indicates that the PU film met the 18 POA design-target with a 

hole diameter of 0.00624 in. The perforate sheet is 34 POA with a hole diameter of 0.05 in. before 

bonding. The installed POA is approximately 32 ± 2 with a hole diameter between 0.048 to 0.050 in.

3.0 Test Program

3.1 Test Site Characteristics

The test site has negligible influence on the engine’s noise generating processes and, except for 

ground surface effects, negligible influence on the propagation of sound from the engine to the 

microphone. See Figure 25.

3.1.1 Overall Site

The C-11 test site is located in an open area having relatively flat terrain which is free of structures 

and other obstacles that affect far-field sound pressure measurements.
NASA/CR—2014-218088 25

Figure 25.—Overall View of PW C11 Stand Far Field Acoustic Arena.

One-third octave band sound pressure levels of ambient noise are sufficiently low so that engine noise 

measurements are not significantly contaminated by noise from environmental or manmade sources.

3.1.2 Engine Support Structure and Test Engine

The engine support structure is designed for minimal sound interference characteristics. That is, the 

structure does not impede sound propagation nor does it have any acoustically reflective surfaces close to 

noise radiating regions. The engine centerline height above the acoustical arena surface is 15 ft (4.57 m), 

or no less than one and one half times the fan diameter.

The test engine was a PW4098. This is a modern high bypass turbofan engine with wide chord 

shroudless fan blades. The 112 in. diameter fan rotor has 22 blades and the Fan Exit Guide Vanes consists 

of 60 static airfoils. There are 10 Intermediate Case struts behinds the FEGV row including the lower duct 

bifurcation. The engine as configured for Phase 1B testing is shown on the test stand with a production 

bill of material inlet, but with a mini-bellmouth lip at the highlight in Figure 26.

Static noise tests in the Fall of 1998 and summer of 1999 evaluated 1) an innovative inlet design 

(Boeing Ideal Scarf inlet) with advanced acoustic treatment, 2) advanced acoustic treatment installed in 

the engine forward fan containment case (FFCC), 3) the impact of inlet probes on fan inlet noise and 4) 

the effect of hard walled sections representative of production type FFCC liner joints.

3.1.3 Acoustical Arena

The acoustical arena is the surface area that extends from beneath the engine to a distance that is at 

least 25 ft (7.6 m) beyond all microphones. The area has the following characteristics: Flat with no 

undulations that cause focusing or scattering of sound or collection of standing water. A slight slope or 

curvature subtending about 0.25° (4 cm/10 m) has been built in to aid in drainage. A thermally reflective 

surface of light concrete is used to minimize solar heating and thermal gradients near the ground. The 

surface is uniformly smooth and hard so as to approximate the acoustical impedance of a perfect sound 

reflector over the frequency range of interest.

3.1.4 Acoustic Barriers

Acoustic barriers were set up to block aft radiated noise from contaminating the inlet radiated noise 

measured by the far field microphones. The walls were about 30 ft (9 m) in height and for this test are 

arranged as shown in Figure 27.
NASA/CR—2014-218088 26

Figure 26.—PW4098 Test Engine with production flight inlet.

Figure 27.—Far field mic array and aft acoustic barriers.
NASA/CR—2014-218088 27

3.2 Engine and Noise Measurement

3.2.1 Data Acquisition and Reduction System

3.2.1.1 Acoustical Data Acquisition and Real Time Analysis System

The acoustical far field data acquisition system consists of microphones, signal conditioning and 

transmission components (e.g., pre-amplifiers, power supplies, and cables) and includes data storage and 

retrieval equipment (e.g., magnetic tape recorders). The objective of this system is to acquire accurate 

measurements of engine sound pressure levels and do on-line digital signal analysis over the range of onethird octave bands having nominal center frequencies from 50 to 10,000 Hz. Deviations from a uniform 

frequency response are determined from instrument manufacturer’s data, calibrations, or combinations of 

both. A Dell GXA computer is programmed to operate two Honeywell 28 track FM tape recorders, 12 

B&K, model 2131, one-third octave band analyzers, six B&K, model 2032, Spectrum Analyzers, and a 

Computer Products Model 7431 Real Time Peripheral (RTP) input/output device. During data 

acquisition, tape recording and one-third octave analysis of up to 70 acoustic signals may be obtained.

Additionally, measurements of engine speeds, temperatures, relative humidity, and wind conditions are 

measured at a rate of nearly two times per second.

Internal inlet noise in the Boeing Ideal Scarf Inlet was recorded on magnetic tape during Phase 1A by 

six Kulites mounted in two axial arrays of three each along the crown and keel. Thirteen static pressure 

taps were installed in the inlet in axial arrays along the crown and at ±90° from top dead center. 

Location of Scarf Inlet Kulite Microphones

Approximate Distance From A Flange (in.) .

Along Keel 3.4, 40.6, 80.3 (K1,K2,K3)

Along Crown 3.4, 29.4, 61.4 (K4,K5,K6)

Location of Scarf Inlet Pressure Taps 

Approximate Distance From A Flange (in.)

Along Crown (0°) 63.6, 58.8, 56.7, 53.2, 50.0, 44.9, 39.9, 34.9 (p11-p18) 

70.9,65.7, 63.6, 60.6, 55.4 (p1-p6, p7-p10)

A special array of 198 dynamic pressure transducers were arrayed on the Inflow Control Device 

(ICD) in Phase 1A of the test. The ICD is a large spherical shaped inflow conditioner mounted in front of 

the engine during acoustic testing. The ICD can be seen in Figure 25. The array was connected to a 

special data acquisition and analysis system for phased array analysis, mode and source identification.

During Phase 1B testing, an internal array of transducers was mounted in the inlet adapter ring

between the inlet and the engine front fan case. The array consisted of 26 Kulites arranged in a 180°

circumferential pattern for rotating mode measurements. Four blade mounted transducers were installed 

on the blade pressure surface near the tip. Special transducer locations are shown in Figure 28. A 

customized P&W telemetry system with a 1 per rev trigger signal was used to transmit data from the 

rotating reference frame to an antenna mounted on the inlet and the signals were transmitted by cable into 

the control room for analysis with a data system supplied by Boeing.

The far field microphone and related components had the following characteristics: Microphones are 

located on a 150 ft (45.7 m) radius array. The microphones are 1/2 in. (1.27 cm) B&K 4134, & Larson 

Davis, 2559 (pressure type), condenser cartridges and installed 1/4 in. (7 mm) above the acoustic arena 

ground plane, with symmetrical protective angular grids on. A diagram of the far field microphone array 

is shown in Figure 27. The microphone angles measured from forward CL and centered on MARP for 

this test were:

Most Configurations: 24 angles, 10° to 120° at 5°, 130°

Configuration 17: 32 angles, 5° to 160° at 5°
NASA/CR—2014-218088 28

Figure 28.—Locations of special internal transducers

The frequency response and stability characteristics were within those specified by the SAE 

Aerospace Recommended Practice, ARP-1846, February 1990. No windscreens were installed. B&K 

preamplifiers Type 2619, or equivalent were used. Cables were routed from the microphone location over 

the surface to central points close by where they were attached to underground cables extending into the 

control room where they attach to the B&K, Type WB0690, Power Supply and Signal Conditioners.

3.2.1.2 Meteorological Measurement System

The meteorological measurement system consisted of instruments to measure wind speed, wind 

direction, ambient air temperature, relative humidity and atmospheric pressure. The objective was to 

monitor meteorological conditions to ensure that static noise tests were conducted within acceptable 

meteorological limits and to provide information needed for acoustical data normalization to reference 

meteorological conditions

Meteorological sensors for wind (speed and direction), air temperature and relative humidity were 

located in the vicinity of the microphone array over the same surface that extends from the engine to the 

far-field microphones. Wind speed and direction were measured at approximately engine centerline 

measure ground wind speed to provide guidance on wind shear effects on noise radiation. Ambient air 

temperature was measured at approximate engine centerline height, 15 ft (4.6 m) and about 1/2 in.

(1.27 cm) above the ground surface.

Relative humidity is calculated from measurements of dry bulb air temperature and temperature of the 

dew point taken at approximately engine centerline height. Ambient atmospheric pressure is measured in 

the vicinity of the acoustical arena.
NASA/CR—2014-218088 29

3.2.1.3 Engine Performance Measurements

Instruments are provided to define the prime engine power setting parameter (e.g., fan rotor speed). 

The objectives are to ensure that the engine is set at the desired operating condition and is stable when 

sound pressure signals are recorded. The instruments have the capability of determining the value of the 

power setting parameter within an accuracy consistent with a rotor speed measurement accuracy of

±1 percent. A once per rev rotor speed signal was provided for the mode measurement analysis and the 

blade mounted transducers.

3.2.2 Acoustical Calibrations

3.2.2.1 Calibrations—General

All components of the test instrumentation were calibrated and certified to comply with the 

manufacturer’s specifications and FAR 36 practices.

3.2.2.2 Calibrations—Frequency Response 

Calibrations were performed to determine deviations of the entire far field recording and reproducing 

and real time data system from a uniform frequency response. The non-uniformity’s were determined in 

terms of time averaged one-third octave band sound pressure levels measured at each preferred one-third 

octave band center frequency reference signal (250 Hz).

The one-third octave band and narrowband corrections required to adjust the recording and reproducing 

and real time data systems to a uniform frequency response were applied to the measured data for 

microphone frequency response, cable losses, analyzer center frequency response and filter shape response.

3.2.2.3 Calibrations—Amplitude

Before and after each test period the overall electro-acoustical response of the microphone system 

was determined by use of a signal from a calibrated acoustical calibrator (pistonphone) generating a 

known sound pressure level, 124.0 dB at 250 Hz. The level was corrected by the difference in output level 

of the pistonphone between the calibrated level at standard day barometric pressure and the output level at 

the barometric pressure when the pistonphone was applied to the microphones.

3.2.3 Test Procedures, Engine and Noise Data Acquisition

3.2.3.1 General Test and Meteorological Conditions

The acoustic surface was free of extraneous clutter, standing water, or other objects that could alter the 

nominal acoustical characteristics of the surface. Testing was conducted within certain atmospheric condition 

limits. Examples of limits for temperature, relative humidity and wind conditions that were used are:

Wind Speed: ± 8 kn average At engine CL height

± 10 kn maximum At engine CL height

Cross Wind: ± 6.5 kn average At engine CL height

Surface Wind: ± 7 kn maximum At 2 in. above ground plane

Temperature 36 to 95 °F ( 2.2 to 35 °C )

Temperature Gradient (Surface Temp.-Centerline Temp. ) ± 7.0 °F (3.9 °C) 

Relative Humidity 20 to 95 percent

Precipitation: None permitted 

3.2.3.1.1 Typical Acoustic Test Data Acquisition Procedures

A single frequency acoustic calibration of the microphone systems was performed with Bruel & Kjaer 

Model 4220 or 4228 pistonphone. The pistonphone output had a current laboratory calibration certificate 
NASA/CR—2014-218088 30

traceable to the NIST. Immediately before each data recording session, the pistonphone calibrator was 

applied to each microphone system. The calibrated output level was normalized to a constant output voltage 

level at the input to the 1/3 octave and narrow band analyzers and the tape recorders and recorded on the 

pistonphone calibration record. Immediately after each recording session, the pistonphone calibration 

procedure was repeated (without normalization adjustments) and the system’s performance was accepted if 

during the acquisition period the sensitivity variation of each microphone system did not exceed ± 0.5 dB

The electronic equivalent sound pressure reference level sinewave at 250 Hz was applied to all 

channels of the real time analyzer system and time averaged for a minimum of 30 sec to establish a sound 

pressure level reference for the analyzers. This was done at the start and end of every data recording 

session. The reference level in volts with respect to the pistonphone calibration level in dB was noted on 

the recorder log sheet. At the beginning and end of every data acquisition period, a pink noise signal, at 

optimum recording level, was applied to and analyzed by the real time analyzers for a minimum of 30 sec. 

The level of the pink noise was set at an optimum recording level. Data were considered reliable if the 

difference between the pink noise at the beginning and end of each acquisition period or each tape did not 

exceed ± 0.8 dB for any of the one-third octave bands.

An Inflow Control Device (ICD), designed to minimize inflow distortion and turbulence into the fan 

was used. It is intended to simulate the clean inflow conditions seen by the engine in flight. It consisted of 

a support structure which includes the frame, transport wheels, perforated plate and honeycomb panels. 

The ICD was positioned such that the engine inlet protruded into the ICD. Both the Boeing Production 

inlet and the Boeing Ideal Scarf inlet were set to similar penetrations into the ICD. A non-porous sailcloth seal was attached to the ICD and the engine to eliminate flow leakage through the annulus between 

the inlet and the ICD opening. The ICD was inspected for damage and screen blockage prior to testing.

Just prior to engine start up for each data recording session, an ambient noise data record was 

acquired. The data were acquired on the real time and the tape recording system. The real time data were 

reviewed immediately prior to engine start up to ensure acceptable ambient noise conditions.

A test matrix of engine speed or power settings was defined prior to conduct of test. The test conditions 

included rotor speeds typical of the acoustic conditions of sideline, cutback and approach. Configuration 

descriptions are provided in Table 4, and pictorial representations are shown in Figure 29. A complete list of 

individual data points by Configuration and Run Number are provided in the Appendix. Selected points were 

repeated once or twice to establish repeatability or confidence limits based on data scatter. Prior to the first 

data acquisition of each data recording session, the engine was temperature stabilized for 5 min. For each test 

condition, before initiating data acquisition, the engine was stabilized such that the engine fan corrected 

rotational speed did not vary more than ± 15 rpm or ± 1 percent of the lowest engine speed recorded.

Real time acoustic data were acquired using an averaging time of 32 sec during which one-third 

octave and narrowband data were processed simultaneously from each microphone. Groups of two 12 

microphone sequences were analyzed sequentially during a data point acquisition. Data from the 24 

microphone array were acquired in about 2 min. Third-octave band data was reviewed on stand within 

minutes of data acquisition and displayed graphically on computer terminals set up in the test stand 

control room. Each data record was identified by a unique record number (see the Appendix). All backup 

acoustic data records on magnetic tape were 60 sec in length. 

The low-pressure shaft rotational speed was sampled at a sufficiently high rate such that a continuous 

representation of this parameter was displayed. The once per rev signal used during Phase 1B testing was 

fed to the Boeing data analysis system for the blade mounted transducers. Temperature, relative humidity, 

wind speed and direction were measured during the period of time when sound pressure signals were 

being measured, recorded and displayed on-line. The speed signal was also recorded on magnetic tape 

with the acoustic signals. The rotational speed during data acquisition was averaged and displayed after 

the data point acquisition. It also displayed the maximum and minimum deviations of engine speed during 

the data acquisition and analysis time.
NASA/CR—2014-218088 31

TABLE 4.—ACOUSTIC CONFIGURATION SUMMARY

Config.

and 

run no

Inlet Wall Front Fan 

Containment 

Case (FFCC)

Inlet treatment Front FCC 

treatment

Inlet 

position

Objective of Test

1

16078

PW4098

Flight Lip

Aft BOM HW HW Keel to mics HW Inlet Production Baseline

2 

16079

PW4098

Flight Lip

Aft BOM Treated BOM Treated Keel to mics Prod Inlet Treatment versus HW

3

16082

Boeing 

Ideal Scarf

Aft Modified 

FCC

HW HW Keel to mics Effect of HW Scarf versus HW Base

12

16083

Boeing 

Ideal Scarf

Aft Modified 

FCC

Treated diffuser, HW 

Lip

HW Keel to mics Effect of inlet treatment forward of the throat 

with HW FFCC

14

16084

Boeing 

Ideal Scarf

Aft Modified 

FCC

Treated diffuser and lip HW Keel to mics Effect of inlet treatment forward of the throat 

with HW FCC

4

16085

Boeing 

Ideal Scarf

Aft Modified 

FCC

Treated diffuser and lip Simulated

BOM

Keel to mics Effect of Scarf Trt versus HW and Trt Scarf 

Versus Trt BOM

5

16086

Boeing 

Ideal Scarf

Aft Modified 

FCC

Treated diffuser and lip Adv 2 in.

Deep Dbl Layer (R1)

Keel to mics Advanced FFCC Treatment R1 

Compare to BOM base FFCC 

6

16087

Boeing 

Ideal Scarf

Aft Modified 

FCC

Treated diffuser and lip Adv 2 in.

Deep Dbl Layer (R2) 

(Tear off Perf. Poly)

Keel to mics Advanced FFCC Treatment R2 

Compare to BOM base FFCC

And Adv R1

7

16088

Boeing 

Ideal Scarf

Aft Modified 

FCC

Treated diffuser and lip Adv 2 in.

Deep Dbl Layer (R2)

Keel to mics Effect of HW FFCC Splices

8 Eq Sp 3-in. wide axial splices

8

16089

Boeing 

Ideal Scarf

Aft Modified 

FCC

Treated diffuser and lip Adv 2 in.

Deep Dbl Layer (R2)

Keel to mics Effect of HW FFCC Splices

3 Eq Sp 3-in. wide axial splices

9

16090

Boeing 

Ideal Scarf

Aft Modified 

FCC

Treated diffuser and lip Adv 2 in.

Deep Dbl Layer (R2)

Keel to mics Effect of HW FFCC Splices

3 Eq Sp 1.25-in. wide axial splices

11

16091

Boeing 

Ideal Scarf

Aft Modified 

FCC

Boeing Ideal Adv 2 in.

Deep Dbl Layer (R2)

Keel to mics Effect of HW FFCC Splices

3 Eq Sp 0.625-in. wide splices

10

16092

Boeing 

Ideal Scarf

Aft Modified 

FCC

Treated diffuser and lip Adv 2 in.

Deep Dbl Layer (R2)

Keel to mics Effect of HW FFCC Splices

3 Eq Sp 1.25-in. axial wide splices + 6 in.

circumferential HW area in front of fan

End of Phase 1A Testing

Start of Phase 1B Testing

P3

16101

PW4098.

BM Lip

Aft BOM BOM Treated BOM treated Keel to mics Diagnostics: Prod inlet baseline with BOM 

Trt FFCC with ICD ground plane

P3

16102

PW4098.

BM Lip

Aft BOM BOM Treated BOM treated Keel to mics Diagnostics: Prod inlet baseline with BOM 

Trt FFCC with NO ICD ground plane

S1

16103

Boeing 

Ideal Scarf

Aft BOM Treated diffuser and lip BOM treated Keel to Mics Diagnostics: BMTs, FFCC Modal array NO 

ICD ground plane

S3

16104

Boeing 

Ideal Scarf

Aft BOM Treated diffuser and lip BOM treated Keel to Mics Diagnostics: BMTs, FFCC Modal array, Flow 

viz tufts near inlet crown, NO ICD ground 

plane

S13A

16105

Boeing 

Ideal Scarf

Aft BOM HW bottom diffuser and 

lip

BOM treated Keel to mics Effect of inlet treatment at keel; HW ±45° 

about keel from "A" flange to inlet highlight

S15A-1

16106

Boeing 

Ideal Scarf

Aft BOM HW circumfer. strip BOM treated Keel to mics Effect of Inlet HW 3-in. wide circumferential 

strip just fwd of "A" flange 

S15A-3

16107

Boeing 

Ideal Scarf

Aft BOM Treated diffuser and lip BOM treated Keel to mics Effect of HW Inlet Splices

3 Eq Sp 3- by 6-in. axial splices just fwd of 

"A" flange

S15-A4

16108

Boeing 

Ideal Scarf

Aft BOM HW throat and lip at 

keel

BOM treated Keel to mics Effect of inlet throat and lip treatment at keel; 

HW ±45° about keel 

S16

16109

Boeing 

Ideal Scarf

Aft BOM Treated diffuser and lip BOM treated Keel rotated 

with keel to 

ground

Effect of scarf geometry on directivity

S13-A1

16112

Boeing 

Ideal Scarf

Aft BOM HW at keel diffuser 

only

BOM treated Keel rotated 

with keel to 

ground

Effect of inlet diffuser treatment at keel; HW 

±45° about keel from "A" flange to inlet 

throat

S17

16113

Boeing 

Ideal Scarf

None BOM Treated diffuser and lip BOM treated Keel To 

Mics

Contribution of aft radiated noise with Boeing 

Ideal Scarf inlet; Compare with Run 16065, 

Production BOM treated inlet
NASA/CR—2014-218088 32

Figure 29.—Phase 1A and 1B Configurations.
NASA/CR—2014-218088 33

A data point information summary and recorder log were maintained for each data acquisition period. 

The log contains pertinent information about the test; especially the engine model, engine number, test 

date, configuration definition, low rotor speed, meteorological conditions, time and date. The same 

information was also part of each data record as stored on the PW Acoustic Data Base and was included 

when the data is transmitted to the customer. See the Appendix. A chronological record of the record 

number, engine low rotor speed, time of data point, outside temperature, relative humidity, and a channel 

by channel record of the amplifier gain or attenuator settings were also maintained.

At the conclusion of the acoustic data acquisition for each engine condition, the data point was accepted 

based on meeting all the following criteria: 1) Data acquisition successfully completed and all required 

acoustical signals appeared good. 2) All weather parameters remained within prescribed limits throughout 

the acquisition time period. 3) The engine performance was within the prescribed target condition limits. 4) 

There was no suspected contamination of the signals from aircraft flyovers or other test stand.

3.2.4 Test Data Processing Procedures for Far Field Real Time Data 

3.2.4.1 Data Corrections

Appropriate corrections to the time averaged one-third octave band and narrowband sound pressure 

levels were applied to account for:

• Microphone pressure response

• Pressure-to-free-field response for 90° incidence, microphone protective grids on 

• Pistonphone reference amplitude response

• Microphone system electrical frequency response

• Real time data reduction system response

○ Analyzer effective bandwidth response

3.2.4.2 Adjustment to Reference Day Atmospheric Conditions

Differences between test and reference day atmospheric absorption coefficients obtained from ARP 

866A were applied to each appropriate one-third octave band sound pressure level over the measurement 

distance. The reference day conditions are 77 °F (25 °C) and 70 percent relative humidity.

4.0 Test Results

4.1 Far Field Data

This section presents the far field noise results for all the configurations tested. All data is inlet radiated 

noise and has been “flown” at typical 777 approach (Mtip ~ 0.8) and cutback (Mtip ~ 1.1) conditions and 

altitudes. The data has been adjusted to remove the “leakage” of low frequency jet noise around the aft 

barrier at high powers. Results and comparisons are presented for the following configurations:

• Production inlet and hardwalled production inlet – baseline

• Hardwalled production inlet and hardwalled scarf inlet

• Benefits of treated scarf inlet plus front fan containment case (FFCC) configurations

• FFCC benefits

• Scarf inlet diffuser treatment benefits

• Inlet Lip Treatment Benefits

• FFCC splice penalties

• Inlet splice penalties

• Scarf inlet orientation effects

• Scarf inlet effect on aft noise
NASA/CR—2014-218088 34

For each of these comparisons a schematic is presented that describes the configurations being 

compared which is followed by a bar graph that shows the comparison benefit or debit. Following these 

results graphs are presented that show noise measurements for approach and cutback versus far field mic 

angle and SPL versus frequency at the mic angle position of interest.

4.1.1 Production Inlet Versus Hardwall Production Inlet—Baseline 

The first result shown in Figure 30 presents a baseline comparison of the fully treated production inlet 

compared to a hardwalled production inlet to establish the benefit for production inlet treatment. As 

indicated in the bar graph the treated inlet provides a 7.2 dB benefit at cutback and a 2.8 dB benefit at 

approach. These attenuations represent noise energy reductions of about 80 and 45 percent, respectively. 

As shown in Figure 31, significant liner attenuations are evident over a wide range of angles and 

frequencies due to the current production liner. 

Figure 30.—Treated production inlet versus hardwall production inlet.

Figure 31.—Production inlet and case noise measurements.
NASA/CR—2014-218088 35

Figure 32.—Hardwall scarf versus hardwall production inlet.

Figure 33.—Hardwall scarf inlet versus hardwall production inlet noise measurements.

4.1.2 Hardwalled Scarf Inlet Versus Hardwalled Production Inlet

In comparing the hardwalled scarf inlet and the hardwalled production inlet, Figure 32 shows that the 

scarf inlet is noisier by 3.0 dB at cutback and 0.7 dB at approach than the production inlet. As shown in 

Figure 33, at approach power the scarf inlet was actually quieter at most angles and frequencies, but an 

inspection of the spectra shows that there is a source of BPF and 2 BPF noise for the scarf inlet that has 

been attributed to flow distortion from the crown region of this inlet. Similarly at cutback power, there is 

an increase in BPF noise and buzzsaw noise for the scarf inlet, particularly at angles near 50°. These 

effects will be discussed in more detail later in this report. 

4.1.3 Benefits of Treated Scarf Inlet Plus FFCC

In Figure 34 the package of principle features of this noise reduction program are compared to a fully 

hardwalled version of the scarf inlet. Configuration 5, which comprises the fully treated scarf inlet and the 

R1 BFG FFCC (two in. deep double layer treatment with perforated polyurethane face sheet) and 

Configuration 6 featuring the fully treated scarf inlet with R2 FFCC (R1 case with the perforated 

polyurethane removed) are compared to a fully hardwalled scarf inlet and FFCC (Configuration 3). The 

fully treated scarf inlet with a simulated production FFCC (Configuration 4) is also compared to the fully 

hardwalled scarf inlet and FFCC.
NASA/CR—2014-218088 36

As indicated in Figure 34, the fully treated scarf inlet with the R1 FFCC produced the greatest noise 

reduction compared to the fully hardwalled inlet and FFCC, indicating that the higher resistance liner 

associated with the perf poly face sheet is better, particularly at cutback power where the combined inlet 

and fan case treatment provided almost 13 EPNdB inlet noise reduction. This corresponds to noise energy 

suppression of about 95 percent. As shown in Figure 35, the extraneous source of BPF and buzzsaw noise 

from the scarf inlet has been attenuated by as much as 25 dB at cutback power.

The bar chart of Figure 36 shows how the best scarf inlet configuration (C5) compares to the fully 

treated production inlet (C1) when both are referenced to the hardwalled production inlet and FFCC. The 

chart shows the C5 scarf inlet configuration provides an 80 percent dB suppression improvement at 

approach and a 50 percent dB improvement at cutback compared to the production inlet (C1).

Figure 34.—Total scarf inlet and FFCC case effectiveness.

Figure 35.—Total scarf inlet and FFCC case noise measurements.
NASA/CR—2014-218088 37

Figure 36.—Best scarf inlet benefit relative to production inlet.

Figure 37.—Forward fan containment case benefits.

4.1.4 Forward Fan Case Treatment Benefits

Figure 37 shows the results of comparisons of the R1, R2, and simulated BOM cases to a hardwalled 

case, all when installed with a fully treated scarf inlet. The current BOM liner in the PW4098 was only 

moderately effective at approach power because of area and backing depth constraints imposed by fan 

containment concerns. At cutback power this BOM treatment actually increases noise relative to the 

hardwalled case. This is attributed, as discussed in more detail later, to noise scattering by the number and 

width of hardwall splices in the BOM case, which separate the three circumferential segments in this 

design. At approach power it can be seen that both the advanced liner impedances provided almost 

1 EPNdB noise suppression (Figure 38). However, at cutback power the R1 resistance associated with the 

perforated poly film provided the best attenuation. Evidently the removal of this film, providing the R2 

impedance, uncovered some impedance non-uniformities that actually increased the noise at cutback 

power through noise scattering. 
NASA/CR—2014-218088 38

Figure 38.—Forward fan containment case treatment noise measurements.

4.1.5 Scarf Inlet Diffuser Treatment Benefits

In order to evaluate the effectiveness of the diffuser section of the scarf inlet, comparisons were made 

between a fully treated diffuser (C12) and a hardwalled diffuser (C3) with the inlet lip and FFCC 

hardwalled. A comparison was also made between a diffuser with its bottom 90° hardwalled (C13A-1) 

and fully treated (C S1) since Boeing analytical studies indicated that bottom diffuser treatment may be 

ineffective. Figure 39 presents these configurations and the corresponding noise differences. The bar 

chart shows that full diffuser treatment relative to hardwall provides a 4.5 dB benefit at approach and a 

11.7 dB benefit at cutback. It also shows that 90° of hardwall on the diffuser bottom produces very little 

noise difference relative to a fully treated diffuser, thereby validating the Boeing analysis. Figure 40 and 

Figure 41 show the noise directivity and spectral features of these benefits.
NASA/CR—2014-218088 39

Figure 39.—Diffuser treatment benefits.

Figure 40.—Full diffuser treatment noise measurements.
NASA/CR—2014-218088 40

Figure 41.—Bottom diffuser treatment noise measurements.

Figure 42.—scarf lip treatment benefits.

4.1.6 Inlet Lip Treatment Benefits

The effectiveness of treatment on the lip section of the inlet was studied by comparing a fully treated 

lip (C14) to a hardwalled lip (C12), both with a hardwalled FFCC and by comparing a fully treated lip 

(C S1) to a lip with the bottom 90° hardwalled (C15A-4), both with the BOM FFCC. The results show 

that full lip treatment is better than a full hardwall lip by 1.5 dB at cutback and ~0.1 dB at approach. A lip 

with the bottom 90° hardwalled is noisier than a fully treated lip by ~0.2 dB at cutback and 0.5 dB at 

approach. Figure 42 and Figure 43 show the results.
NASA/CR—2014-218088 41

Figure 43.—Full lip treatment with hardwall case noise measurements.

4.1.7 Fan Case Liner Splice Penalties

Configurations were tested to determine what effect treatment splices in the FFCC have on inlet 

noise. A number of previous studies at PW and by other engine and nacelle manufacturers have indicated 

that these splices can have a large impact on perceived noise levels, primarily due to increases in the 

blade passing frequency (BPF) tone levels. This is particularly the case at high subsonic and low 

supersonic fan tip speeds. PW has developed an analytical model for this phenomenon that involves 

scattering of BPF noise by the impedance discontinuities that these splices represent. This scattering 

involves a transfer of energy from the very high levels of axially decaying BPF tone energy very near the 

fan into propagating modes at the BPF frequency. The degree of energy scattering is predicted to increase 

with both the splice width and the splice number. If the treatment can be made very uniform near the fan, 

this high amplitude decaying sound field will decay uninterrupted and will not scatter high sound energy 

into propagating modes. In many cases this is not possible and significant scattering has been observed. In 

some cases, it has even been found desirable to completely hardwall the treated segments just forward of 

the fan in order to avoid this BPF noise increase. 

In the current test program the number and the width of various simulated splices (simulated by 

applying strips of tape for testing) were investigated as shown in Figure 44 to Figure 47. The results show 

that uniformity of treatment near the fan is very important. Figure 44 shows the effects of splice number, 

for a constant splice width of 3 in. As expected eight splices results in more scattering (higher noise) than 

three splices at both approach and cutback conditions. At cutback power, where the blade tip speeds are 

slightly supersonic, the BPF noise energy is carried by a series of oblique shock waves from the tip 

regions of each blade. The strength of these shockwaves decays rapidly forward of the fan blades in a 

hardwall duct. But if impedance discontinuities (e.g., splices) are placed in this region, the shockwave 

energy is scattered into propagating modes, whose levels can be quite high, particularly if the number or 

width of these discontinuities is large. Figure 44 shows that at this condition (i.e., cutback power), a large 

penalty of 4 EPNdB was measured for C7, featuring eight splices, 3 in. wide. Figure 45 shows that this is 

due primarily to the fact that BPF levels were increased by almost 10 dB for this configuration, relative to 

a uniform hardwalled FFCC. 
NASA/CR—2014-218088 42

Figure 44.—Effect of splice number.

Figure 45.—Splice number noise measurements.

The effect of splice widths ranging from 0.62-in. wide to 3.0 in. wide for a constant splice number 

(i.e., three) is shown in Figure 46 and Figure 47. Here it can be seen that there is very little penalty if the 

splices can be made thin enough (e.g., about 0.62-in. wide). However, when the three splices were made 

to be 3 in. wide, a penalty of up to 1.6 EPNdB resulted at cutback power. Again, this resulted from a BPF 

tone level increase of about 6 dB, as shown in Figure 47. 
NASA/CR—2014-218088 43

Figure 46.—Fan case splice penalties.

Figure 47.—Fan splice width noise measurements.
NASA/CR—2014-218088 44

4.1.8 Inlet Liner Splice Penalties

The effects of placing splices in the inlet liner forward of the FFCC was also measured. A 

circumferential splice forward of “A” flange (C15A-1) and three 3-in. wide axial splices (C15A-3) were 

investigated. Comparisons were made to C S1and are shown in Figure 48 to Figure 52. The effect of the 

uniform circumferential splice is seen to be about 0.2 to 0.4 EPNdB. This increase is entirely consistent 

with the loss of treated area associated with configuration 15A-1. 

The effect of three, 3 in. splices in the inlet are shown in Figure 50. The noise penalty associated with 

these splices is seen to be minimal. For reference, the much larger penalties for this same splice 

simulation in the FFCC is also shown. The reason for the difference is that the inlet splices are spaced 

axially at a much larger distance from the fan blades, thus allowing the high levels of decaying BPF tone 

levels to decay naturally before getting scattered by the inlet splices. The message here is that it is 

extremely important to have treatment uniformity near the fan, but it is less important that uniformity 

exists for treatment further upstream. 

Figure 48.—Inlet circumferential splice penalty.

Figure 49.—Inlet circumferential splice noise measurements.
NASA/CR—2014-218088 45

Figure 50.—Inlet versus fan case axial splice penalties.

Figure 51.—Inlet orientation effects.

4.1.9 Inlet Orientation Effects

The effects of inlet orientation—scarf keel pointed toward the ground (C16) versus scarf keel pointed 

towards the mics (C S1)—were measured. The results showed no apparent benefit for the scarf keel being 

oriented towards the mics. As shown in Figure 50, noise measurements with the keel pointed to the mics 

were actually higher by 1.0 EPNdB at approach and ~0.1 EPNdB at cutback.

4.1.10 Scarf Effect on Aft Noise

The effect of the scarf inlet on aft fan noise was measured by removing the wall barriers and 

comparing a production inlet configuration with the scarf inlet and a BOM FFCC, (C17). The results in 

Figure 53 show that aft fan noise is not appreciably influenced by the scarf inlet.
NASA/CR—2014-218088 46

Figure 52.—Inlet orientation effects: benefits of scarf toward mics.

Figure 53.—Scarf effect on aft noise.

Figure 54.—Scarf effects on aft noise—noise measurements.
NASA/CR—2014-218088 47

4.1.11 Summary of Various Treatment Benefits

Figure 55(a) and (b) summarize the benefits at cutback and approach powers, respectively, associated 

with various segments of treatment that were tested in the production inlet, the scarf inlet and the forward 

fan containment case (FFCC). These benefits are plotted against treatment length divided by duct radius.

At cutback power, the complete production package (inlet and FFCC) are seen to result in about 7 EPNdB 

noise reduction. The complete scarf inlet (diffuser plus lip), combined with the advanced R1 fan case, 

resulted in a benefit of almost 11 EPNdB, or about 4 EPNdB more than the production design. The 

treatment non-uniformities contained in the R2 and production fan cases actually had a negative benefit 

on noise for reasons discussed earlier. 

At approach power (see Figure 55(b)) the total noise benefit of the scarf inlet and advanced R2 FFCC 

treatment is seen to be about 5 EPNdB, compared to less than 3 EPNdB provided by the production inlet 

and FFCC treatment. The slopes of the attenuation versus L/R curve segments shown in Figure 55(a) and 

(b) are an indication of how effective (e.g., attenuation/treatment area) the various sections of liner are in 

the inlet and FFCC. These slopes are plotted in Figure 56. Here it can be seen that the scarf inlet treatment

is more effective than the production inlet at both approach and cutback conditions. In addition, the R1 

and R2 FFCC liners both provided about 4 EPNdB per L/R at approach power, similar to the scarf inlet. 

Figure 55.—(a) Summary of inlet and fan case treatment benefits at cutback power. 

(b) Summary of inlet and fan case treatment benefits at approach power.
NASA/CR—2014-218088 48

Figure 56.—Relative effectiveness of various liners—normalized by L/R

4.1.12 Conclusions—Far Field Data

• The treated scarf inlet with the advanced (R1) FFCC is 2 to 4 EPNdB quieter in inlet fan noise 

than the production inlet and fan case at approach and cutback powers. Approximately 80 percent

dB nacelle suppression improvement was observed at approach power and 50 percent dB 

suppression improvement at cutback power.

• The Scarf inlet with advanced FFCC has a significant BPF distortion tone, preventing larger 

benefits. Without the extra BPF noise an additional 1 EPNdB in noise reduction was expected.

• The hardwalled scarf inlet has higher buzzsaw noise levels than the production inlet but treatment 

suppressed it very well.

• The bottom 90° of diffuser liner is ineffective at all conditions.

• The scarf lip liner effectiveness at cutback power is about the same per unit length as the diffuser 

treatment.

• The scarf lip liner at approach power is effective over a wide frequency range, but not for the 

BPF distortion tone thus preventing a significant EPNdB benefit.

• At approach power both advanced FFCC liners (R1and R2) have the same effectiveness per unit 

length as the scarf inlet liner.

• At cutback power the higher resistance fan case liner (R1) showed a significant benefit while the 

lower resistance liner (R2) was detrimental, primarily due to impedance non-uniformities.

• Treatment uniformity near the fan is very important. Simulated production type hardwall splices 

in the FFCC liner show large penalties of up to 4 EPNdB at cutback power for eight 3 in. splices. 

Penalties increase with both splice number and width. The mechanism is scattering, by 

impedance discontinuities, of high level but fast decaying blade pressure fields at BPF (m = B) 

into high level propagating modes.

• The treatment uniformity further forward in the inlet is not critical. Penalties associated with axial 

or circumferential splices are associated only with lost treated area.

• There is no apparent benefit of the scarf keel being oriented toward the microphones, relative to 

being oriented 90° away. This may be due to the effect of the scarf geometry on distortion tone 

generation or on azimuthal buzzsaw noise variations.

• Aft fan noise is not influenced appreciably by the scarf inlet. 
NASA/CR—2014-218088 49

4.2 RDIFF Prediction Comparisons

4.2.1 RDIFF Prediction Code 

The RDIFF computer code (Ref. 1) used for the acoustic design of the scarf inlet is a ray tracing code 

in which it was assumed that the noise source consisted of uncorrelated monopoles distributed over the 

outer 1/3 radius of the fan. The full 3D shape of the inlet was modeled with the lining distribution 

accounted for by specifying the lining’s locally reacting impedance in lining segments. Twenty-nine 

lining segments were used to model the inlet and forward fan case. The effect of flow is considered for 

calculation of the liner impedance and for calculation of the ray incidence angle to the lining. The (plane 

wave) reflection coefficient is then calculated for each source ray. No nonlinear propagation effects 

(buzzsaw) are evaluated and ray diffraction is applied at inlet lip to account for sound propagation to the 

rear of the inlet.

4.2.2 Approach Power

4.2.2.1 Production Inlet—Approach Power

Figure 57 compares the RDIFF predictions for the fully treated production inlet (PTr) with test 

results. The predictions and data have been extrapolated to a nominal 777 approach condition. The 

prediction curves were generated by subtracting the production inlet RDIFF predicted 1/3 octave spectral 

and directional 1/3 octave SPL differences, (hardwall – full treatment), from the measured production 

inlet hardwall data. Comparisons are shown for the PNLT time history and spectra at angles from the inlet 

axis of 30° to 80°. The agreement of the prediction curves with the measured data is quite good until 

about 70°.

Figure 57.—Fully treated production inlet noise prediction—approach.
NASA/CR—2014-218088 50

4.2.2.2 Scarf Inlet—Approach Power

Figure 58 shows the same comparison for the hardwall scarf inlet (SHW) at approach power. Again 

the RDIFF predicted difference between the hardwall production inlet (PHW) and the inlet configuration 

of interest was subtracted from the hardwall production inlet test data to obtain the prediction curves. For 

this comparison, only the effect of the inlet shape difference is calculated, no lining effects were included.

It is seen that the prediction is quite good forward of 70°, except for the BPF and 2BPF tone increase 

observed with the scarf inlet relative to the production inlet.

The prediction/data comparison for the fully treated scarf inlet (STr) is shown in Figure 59. Again the 

prediction reference is the fully hardwall production inlet. As for the hardwall scarf inlet, the BPF/2BPF 

tone levels are not predicted by this RDIFF process. Forward of 70°, the predicted and measured spectra 

compare well except for the BPF/2BPF tones. But there are still problems with the aft angles. 

Figure 58.—Hardwall scarf inlet noise prediction—approach.
NASA/CR—2014-218088 51

Figure 59.—Fully treated scarf inlet noise prediction—approach.

Figure 60 shows attenuation directivity examples for three one third octaves centered at 800 Hz, 2 and 

4 kHz. The 800 Hz band contains the BPF, so the comparisons with the measured scarf inlet attenuations 

relative to the production inlet are poor because of the anomalous BPF source that occurred with the scarf 

inlet. Forward of 70°, with the above exception, the RDIFF predicted deltas compare well with the 

measured data. The 2 and 4 kHz comparisons are repeated in Figure 61 with the test results on the same 

plot as the RDIFF predictions for easier comparison. Aft of 70° the scarf inlet lining attenuation (SHWSTr) is somewhat under predicted. The production inlet lining attenuation (PHW-PTr) and the scarf shape 

effect (PHW-SHW) deltas are over predicted around 70° and 80° and mostly under predicted further aft.

In this region the ray diffraction calculation controls the prediction for the scarf inlet but may not control 

the 70° and 80° region for the production inlet (This needs further investigation). The scarf inlet lining 

attenuation includes the effect of a treated lip. It appears that the attenuation of the diffracted rays with lip 

treatment is under predicted and the hardwall lip diffraction is inconsistently predicted.

Predicted attenuation spectra for angles 40°, 60° and 80° are compared to test data in Figure 62. The 

40° and 60° comparisons are repeated in Figure 63 with the test results directly compared to the RDIFF 

predictions for easier comparison. The attenuations relative to the production inlet for the 1/3 octaves in 

the range of the BPF and 2BPF are generally poorly predicted. The measured scarf lining attenuation 

(SHW-STr) is seen to be lower at BPF than the adjacent frequencies as well. This is not predicted by the 

RDIFF code, but this may be related to the mode structure of the anomalous BPF source that occurred 

with the scarf inlet. The BPF tone attenuation for the production inlet lining is reasonably well predicted 

at 60° but is underpredicted at 40°.
NASA/CR—2014-218088 52

Figure 60.—Attenuation directivity—approach power.

Figure 61.—Attenuation directivity—approach power.
NASA/CR—2014-218088 53

Figure 62.—Attenuation spectra—approach power.

Figure 63.—Attenuation spectra—approach power.
NASA/CR—2014-218088 54

4.2.2.2.1 Attenuation Versus Lining Area—Approach Power

Figure 64 compares measured and RDIFF predicted approach inlet noise EPLN attenuations relative 

to the hardwall production inlet EPNL. As shown above, the RDIFF calculation for the scarf inlet 

excludes the effect of the anomalous BPF and 2 BPF tone increases. This plot clearly brings out the 

nearly factor of two lining area increase in the fully treated scarf inlet relative to the fully treated 

production inlet. The RDIFF prediction curve for the production inlet was extended beyond the 

production inlet lining area by first assuming the inlet lip was treated similarly to the fully treated scarf 

inlet by adding lip lining and then “stretching” the inlet so that the lining area equals the lining area of the 

scarf inlet. At the fully treated scarf inlet lining area it is seen that the measured scarf data is about 

1 EPNdB lower than the calculated production inlet level and the “no anomalous tone” scarf prediction is 

about 3 EPNdB quieter. A prediction for the fully treated scarf inlet is also shown assuming the 

production double layer lining in the scarf inlet, rather than the tested triple layer lining. With equal 

linings the “no anomalous tone” scarf prediction is about 2 EPNdB quieter than the “stretched” 

production inlet, essentially maintaining the hardwall scarf inlet shape benefit predicted.

4.2.3 Cutback Power

4.2.3.1 Hardwall Scarf Inlet—Cutback Power

The RDIFF prediction for the hardwall scarf inlet at cutback power is compared with the cutback 

extrapolation of the test data in Figure 65. As for the approach power comparisons, the RDIFF scarf inlet 

hardwall (SHW) prediction curves were calculated by subtracting the RDIFF delta (PHW-SHW) from the 

production hardwall (PHW) test data. Figure 65 shows that the measured cutback power noise level of the 

SHW was significantly higher than the PHW. Examination of narrow band data shows that this increased 

noise is primarily due to buzz saw noise. The increased buzz saw noise cannot be predicted by the RDIFF 

code since it does not predict fan noise source changes or non-linear propagation effects. The RDIFF code 

will predict a directivity change due to the inlet shape but it is clear that in this case the radiated sound 

power is increased. Using the PHW data as a base for the resulting RDIFF prediction, SHW = PHWRDIFF(PHW-SHW) is therefore seen to significantly under predict the measured SHW data.

Figure 64.—Relative attenuation versus lining area—approach power.
NASA/CR—2014-218088 55

Figure 65.—Hardwall Scarf Inlet Noise Prediction—Cutback

4.2.3.2 Fully Treated Scarf Inlet—Cutback Power

Since RDIFF does not predict buzzsaw noise attenuation, it was not possible to use the PHW data as a 

base for the STr predictions at cutback. As an experiment it was decided to try a prediction using the PTr 

data as a base. The results of this calculation are shown in Figure 66. Surprisingly the resulting spectral 

predictions compare fairly well with the measured STr data below frequency 1/3 octave band 

31(1250 Hz). This was especially surprising in light of the large buzz saw noise increase measured for the 

SWH relative to the PHW. For the treated scarf inlet the acoustic lining appears to suppress most of this 

increased buzz saw noise and the difference between the production and scarf acoustically treated inlets 

can be accounted for by linear acoustics. For frequency bands above band 31 the RDIFF treated scarf inlet 

prediction shows much larger suppression relative to the PTr levels than the measured data. This is 

primarily responsible for the large under prediction of the PNLT’s, since this part of the noise spectrum 

has the largest NOY weighting.

Figure 67 shows that the attenuation spectra predicted by RDIFF are fairly similar for approach and 

cutback power conditions. The measured attenuation spectra however are seen to be quite different. For 

frequencies greater than 1200 Hz the measured data attenuations drop off monotonically as engine power 

is increased, contrary to the predictions. The reason for this discrepancy is not known. It is speculated that 

the increasing lip wall Mach numbers with increased engine power may be related. Examination of the 

hardwall lip data showed the same behavior so it was concluded that the attenuation fall off was not due 

to noise generation by the flow over the lip lining. Another possibility is a propagation effect through the 

high Mach number flow, which reduces the hardwall and treated inlet noise levels increasingly as engine 

power is increased while reducing the effect of the lining since the high Mach number flow is adjacent to 

the lining. The other feature of the measured data attenuation spectra seen in Figure 67 worth noting is the 

large low frequency attenuation seen at the cutback condition. This is buzz saw noise attenuation. This is 

clearly not seen in the RDIFF predicted cutback power attenuation spectrum.
NASA/CR—2014-218088 56

Figure 66.—Fully treated scarf inlet noise prediction—cutback.

Figure 67.—Predicted versus measured attenuation spectra.
NASA/CR—2014-218088 57

Figure 68.—Attenuation spectra—cutback power.

Additional cutback power measured and treated attenuation spectra are shown in Figure 68. The 

effects of the measured increased buzz saw noise for the hardwall inlets and the large buzz saw noise 

attenuation for the treated inlets are again contrasted with the RDIFF predictions.

4.2.3.3 Attenuation Versus Lining Area—Cutback Power

Figure 69 is a plot of the scarf inlet EPNL attenuation relative to the production inlet versus lining 

area for cutback power, similar to Figure 64 for approach power. A major difference is that the measured 

treated production inlet data is used as the reference for the treated scarf inlet predictions and the 

increased lining area production inlet predictions because of the inability of RDIFF to predict buzz saw 

noise attenuation. The RDIFF predictions therefore assume no buzz saw noise source increase and no fall 

off of high frequency attenuation as observed in the test data. This is considered a prediction of the noise 

changes relative to the production inlet that would be observed in flight assuming that the high frequency 

attenuation loss and the increased hardwall buzz saw noise are due to static test flow anomalies. As was 

the case for approach power, the prediction for the scarf inlet shows a benefit even if the production inlet 

were “stretched” and lining added to the lip so that they have equal lining area. The benefit of the triple 

layer lining is predicted to be significantly less than was the case for approach power.
NASA/CR—2014-218088 58

Figure 69.—Relative attenuation versus lining area—cutback power.

4.2.4 Scarf Lining Variations

The scarf inlet was tested with a number of lining variations. Figure 70 shows comparisons of the 

measured and RDIFF predicted inlet noise component EPNL attenuations for these variations as follows:

Full Inlet + R1 Case (Scarf full hardwall—Scarf full treatment with the advanced fan case with PU 

overlay (R1)).

Full Diff (Scarf full hardwall—Scarf with hardwall forward fan case and lip, full diffuser 

treatment.

Bott Diff (Scarf with hardwall bottom diffuser, full treatment otherwise—Scarf with full 

treatment)

BOM Case (Scarf with hardwall fan case, full treatment otherwise—Scarf with full 

treatment and simulated BOM fan case)

R1 Case (Scarf with hardwall fan case, full treatment otherwise—Scarf with full 

treatment and R1 fan case)

R2 Case (Scarf with hardwall fan case, full treatment otherwise—Scarf with full 

treatment and simulated R2 fan case)

Full Lip (Scarf with hardwall fan case and lip, treated diffuser—Scarf with hardwall fan 

case, treated lip and treated diffuser)

Bot Lip (Scarf with hardwall lower lip, full treatment otherwise—Scarf with full 

treatment)

“A” Flange (Scarf with 360° hardwall region at “A” flange extending 3 in. upstream into 

inlet—Scarf with full treatment)

As discussed above, the RDIFF code does not predict the buzz saw noise attenuation well. This 

accounts for the poor agreement between the prediction and measurement for the cutback condition for 

the full inlet and full diffuser treatment configurations. The BOM and R2 fan case treatments are seen to 

have increased noise at the cutback condition (because of the hardwall circumferential discontinuities).

The effect of scattering of the rotor bound field by lining discontinuities is not included in the RDIFF 

prediction and therefore results in poor agreement for these cases. The full lip treatment was more.
NASA/CR—2014-218088 59

Figure 70.—Inlet lining elements attenuation.

effective at the cutback condition than predicted. This also appears to be the result of more buzz saw 

attenuation than predicted by RDIFF. At approach power the lip treatment was effective at frequencies 

other than the BPF but the lack of BPF attenuation resulted in no EPNL attenuation. It is unclear why the 

bottom lip liner shows more effectiveness than the entire lip liner at approach power. It is possible that 

hardwalling the upper lip reduced the distortion into the fan which was causing the increased BPF with 

the scarf inlet. Hardwalling approximately 7 ft2 of acoustic lining at the “A” flange was not predicted to 

have any effect on the inlet EPNL. The test results showed a small effect of about 0.5 EPNdB at approach 

and less at cutback.

4.2.5 Conclusions

At the approach power condition it is concluded that the RDIFF predictions for tone and broadband 

lining attenuation generally were within about 15 percent of the measured data in the direct radiation field 

but only within about 50 percent in diffraction dominated field (aft of 70°). However, the very high 

frequency lining attenuation (8 to 10 kHz) was not accurately predicted. The scarf shape shielding effects 

were predicted well except for in the diffraction-dominated field also. The RDIFF prediction process of 

applying the RDIFF predicted noise change (PHW-configuration) to the measured production inlet test 

data was used to estimate the scarf inlet noise if the BPF and 2BPF tone increase did not occur. This is 

believed to be a valid flight condition prediction since it is believed that the flow distortion causing the 

tone increases will not be present with forward motion. It was concluded that without the BPF and 2BPF 

aggravation the 13° biplanar scarf shape is worth about 2 EPNdB and the triple layer lining is worth about 

1 EPNdB relative to a production inlet with lip lining and “stretched” to have the same total lining area as 

in the scarf inlet.

At cutback power it was concluded that the scarf shape significantly increases buzzsaw noise, but, 

except for the high frequencies (2 kHz and higher), the lining attenuated it to levels expected based on the 

measured treated production inlet levels. The RDIFF code however does not predict the attenuation of the 
NASA/CR—2014-218088 60

buzz saw noise. The RDIFF calculated acoustic wave attenuation for the treated scarf inlet relative to the 

treated production inlet at the dominant buzz saw noise frequencies (<1 kHz) appeared to be reasonably 

close to the measured data. Besides buzz saw noise attenuation prediction, the other major miss by the 

RDIFF code was the prediction of the high frequency (2 kHz and higher) noise attenuation at the higher 

power conditions. Contrary to the RDIFF predictions, the measured high frequency attenuations were 

observed to fall off rapidly with increased power setting. It is speculated that this is due to a propagation 

effect through the high Mach number flow, which reduces the hardwall and treated inlet noise levels 

increasingly as engine power is increased while reducing the effect of the lining which is adjacent to the 

high Mach number flow. If this is true the RDIFF high frequency predictions may be valid for flight 

where the lip Mach numbers are much lower.

If the predicted high frequency attenuation is realized, the 13° biplanar scarf shape is worth about 

3 EPNdB and the triple layer lining is worth about 0.75 EPNdB relative to a production inlet with a 

treated lip and “stretched” to have lining area equal to the tested scarf inlet.

4.3 ICD Microphone Array Test Results

4.3.1 Background

As part of the NASA Engine Validation of Noise Reduction Concepts test program, a scarf inlet was 

tested on a PW4098 engine at the Pratt & Whitney static engine test stand at West Palm Beach, Florida in 

November 1998. In an effort to characterize the noise radiation properties of the inlet/engine combination, 

the inflow control device was fitted with 198 Kulite microphones. The Boeing Phased Array Data system 

was used to measure acoustic data from the ICD array. The system was installed at the C-11 stand in 

August 1998 and checked out on a PW4098 certification test prior to the NASA test in November. Phased 

array data was processed to form cross-spectral matrices. These were used in beamforming to make 

images of the fan and low pressure compressor sources. A spinning mode decomposition process was 

applied to create mode plots. Results from the NASA test are presented here for eleven configurations: a 

production inlet with production acoustic lining, the production inlet with the acoustic lining covered with 

tape, and ten configurations of the scarf inlet with various acoustic lining arrangement.

4.3.2 Instrumentation

Kulite model MIC-093 transducers with B-screens were installed nylon sleeves and attached to the 

perforated plate surface of the ICD with screws (Figure 71 to Figure 73). The layout of the transducers 

(Figure 74) was designed for two purposes: beamforming to make images of the noise source distribution 

and spinning mode decomposition. 

Figure 71.—Inflow control device. (Engine depicted is not the PW4098 engine used in the present test.)
NASA/CR—2014-218088 61

Figure 72.—MIC-093 Kulite and the nylon sleeve for mounting it the IDC. The 

signal lead and the vent tube can be seen. A small hook is used to clamp the 

fixture the perforated plate.

Figure 73.—Photograph of the installed microphones.
NASA/CR—2014-218088 62

Figure 74.—Layout of the Kulites on the ICD.

The theoretical performance of the array for separating spinning modes is indicated in Figure 75. 

Paper templates were used to place the microphones near their intended locations (Figure 76).
NASA/CR—2014-218088 63

Figure 75.—Theoretical performance of the array design for spinning 

mode measurements.

Figure 76.—Paper templates used to locate 

microphones on the ICD.

Photogrammetry was used to measure the positions of the Kulites (Table 5).
NASA/CR—2014-218088 64

TABLE 5.—POSITIONS OF THE KULITE MICROPHONES AS DETERMINED BY PHOTOGRAMMETRY

Kulite X Y Z Kulite X Y Z Kulite X Y Z

1

2

3

4

5

6

-15.22

-15.56

-15.25

-15.41

-14.55

-14.63

0.00

-11.36

-30.27

-48.30

-71.95

-96.82

-124.32

-125.90

-122.67

-116.59

-102.45

-79.16

67

68

69

70

71

72

-89.05

-86.95

-86.27

-87.00

-86.14

-86.75

128.99

99.45

69.88

45.04

25.77

11.36

-63.41

-103.52

-124.64

-136.30

-140.92

-142.47

133

134

135

136

137

138

-179.27

-179.27

-179.27

-179.93

-179.27

-179.27

-90.75

-109.19

-92.38

-0.20

92.38

109.19

-62.11

-13.08

59.66

110.40

59.66

-13.08

7

8

9

10

11

12

-16.13

-15.22

-15.20

-14.89

-15.22

-15.22

-118.30

-122.84

-88.92

0.08

89.22

122.84

-41.72

19.14

86.62

124.78

86.57

19.14

73

74

75

76

77

78

-112.47

-114.12

-111.93

-111.88

-111.93

-111.45

0.25

-9.35

-22.45

-41.44

-65.42

-94.46

-140.94

-140.57

-139.56

-134.74

-125.32

-104.46

139

140

141

142

143

144

-178.49

-178.79

-175.85

-178.47

-180.93

-179.86

90.55

65.09

44.57

28.16

14.53

6.04

-61.46

-87.64

-102.10

-108.47

-111.95

-110.17

13

14

15

16

17

18

-15.22

-15.22

-15.19

-14.83

-14.67

-15.65

118.04

96.40

69.27

47.07

28.31

11.56

-39.00

-78.50

-104.58

-116.57

-122.55

-124.52

79

80

81

82

83

84

-111.93

-111.70

-111.93

-111.50

-111.99

-111.93

-124.81

-141.12

-112.09

-0.14

111.46

141.37

-66.39

-1.22

86.15

141.78

86.13

-1.15

145

146

147

148

149

150

-196.76

-196.69

-197.06

-196.64

-196.44

-196.66

-0.39

-5.19

-13.06

-22.43

-35.48

-52.72

-92.27

-91.95

-91.33

-89.85

-85.58

-76.82

19

20

21

22

23

24

-39.60

-37.90

-37.21

-37.79

-36.80

-38.31

2.17

-12.13

-27.59

-48.26

-73.16

-101.35

-134.09

-134.35

-132.43

-126.21

-113.69

-90.47

85

86

87

88

89

90

-113.60

-110.51

-111.45

-112.76

-111.93

-112.21

124.42

94.91

65.58

41.73

23.31

10.32

-67.83

-104.67

-124.72

-134.63

-139.44

-140.02

151

152

153

154

155

156

-197.03

-196.97

-196.97

-196.97

-196.97

-196.66

-73.58

-91.18

-78.81

0.00

78.81

92.17

-55.42

-14.17

47.99

92.27

47.99

-12.64

25

26

27

28

29

30

-38.58

-37.90

-37.80

-37.86

-37.66

-37.90

-125.50

-134.04

-99.94

0.15

99.52

134.04

-49.02

15.12

91.37

134.64

91.28

15.12

91

92

93

94

95

96

-136.96

-137.07

-137.23

-136.08

-136.35

-135.58

-0.01

-8.47

-20.13

-37.23

-58.96

-86.84

-134.96

-134.29

-133.00

-129.48

-121.13

-103.58

157

158

159

160

161

162

-196.46

-196.59

-196.55

-197.67

-198.83

-196.76

74.18

52.96

35.20

21.82

10.80

4.37

-54.12

-75.41

-85.82

-90.38

-91.65

-92.29

31

32

33

34

35

36

-37.90

-36.87

-37.33

-36.96

-37.13

-37.43

126.06

100.32

72.94

47.89

27.53

12.42

-48.01

-91.81

-113.97

-126.35

-132.78

-134.30

97

98

99

100

101

102

-136.10

-135.59

-136.10

-136.10

-136.64

-136.10

-116.56

-134.64

-109.19

0.50

109.45

134.75

-67.90

-6.62

79.20

135.10

79.17

-6.23

163

164

165

166

167

168

-211.16

-211.15

-211.46

-211.79

-212.24

-211.32

0.01

-3.44

-8.74

-15.46

-25.75

-39.31

-71.82

-71.83

-71.36

-69.73

-67.03

-60.05

37

38

39

40

41

42

-62.95

-61.68

-61.74

-63.28

-61.77

-60.31

0.19

-12.09

-27.41

-47.84

-72.60

-103.17

-140.65

-140.16

-138.43

-132.82

-120.50

-96.99

103

104

105

106

107

108

-135.81

-135.99

-136.24

-136.17

-136.10

-137.52

115.67

85.86

58.78

37.31

20.61

8.78

-69.11

-103.55

-121.28

-129.97

-133.31

-134.38

169

170

171

172

173

174

-211.67

-211.71

-211.32

-211.32

-213.41

-211.20

-56.38

-70.41

-62.25

0.00

61.01

70.50

-45.70

-13.86

35.73

71.78

37.86

-12.67

43

44

45

46

47

48

-64.07

-62.07

-61.53

-61.81

-61.66

-62.07

-130.41

-141.01

-107.55

-0.06

107.15

141.01

-55.59

10.05

93.53

140.93

92.51

10.05

109

110

111

112

113

114

-159.25

-159.48

-159.86

-158.44

-159.37

-158.81

0.26

-6.69

-17.26

-32.16

-51.89

-76.67

-124.64

-124.13

-123.30

-120.63

-113.59

-98.18

175

176

177

178

179

180

-211.89

-210.71

-211.33

-211.62

-211.33

-211.52

55.83

39.89

25.78

15.62

7.78

3.02

-44.08

-59.39

-67.76

-70.70

-71.69

-72.21
NASA/CR—2014-218088 65

TABLE 5.—POSITIONS OF THE KULITE MICROPHONES AS DETERMINED BY PHOTOGRAMMETRY

Kulite X Y Z Kulite X Y Z Kulite X Y Z

49

50

51

52

53

54

-62.07

-62.07

-61.19

-62.07

-61.16

-60.07

129.82

102.08

72.81

47.36

27.61

11.09

-55.98

-97.81

-121.62

-133.20

-138.76

-140.97

115

116

117

118

119

120

-158.78

-158.86

-158.78

-158.33

-158.78

-158.78

-105.03

-123.74

-102.59

0.33

102.59

123.89

-66.52

-8.95

70.22

124.09

70.22

-10.34

181

182

183

184

185

186

-223.15

-223.01

-222.68

-222.30

-221.15

-221.89

0.89

-2.20

-4.81

-9.47

-16.95

-25.80

-50.58

-50.85

-50.62

-49.77

-46.05

-41.77

55

56

57

58

59

60

-88.03

-87.16

-89.62

-87.44

-87.01

-86.23

-0.89

-10.50

-24.13

-44.88

-69.91

-99.72

-143.11

-142.36

-139.73

-135.45

-124.69

-102.96

121

122

123

124

125

126

-158.12

-158.29

-158.52

-157.94

-161.76

-158.95

105.25

76.90

51.66

32.11

18.06

5.85

-65.14

-97.06

-112.64

-120.05

-122.58

-125.15

187

188

189

190

191

192

-222.08

-222.23

-221.47

-219.84

-223.67

-222.49

-36.62

-48.10

-42.63

0.37

43.73

48.72

-33.20

-10.01

22.69

48.86

24.44

-9.77

61

62

63

64

65

66

-87.00

-87.15

-83.67

-87.88

-86.71

-87.00

-129.34

-144.26

-110.93

-0.87

111.33

143.48

-62.28

5.52

91.65

143.46

90.17

4.45

127

128

129

130

131

132

-179.83

-179.66

-179.24

-178.66

-179.02

-179.23

-0.06

-6.10

-15.14

-27.45

-44.26

-66.05

-110.11

-109.50

-108.55

-106.25

-100.76

-88.71

193

194

195

196

197

198

-221.95

-221.81

-221.71

-222.13

-222.37

-222.85

37.98

26.33

16.77

10.37

5.89

2.85

-31.41

-41.99

-46.99

-49.14

-49.80

-50.42

Figure 77.—Block diagram of the Boeing Phased 

Array Data Acquisition System.

A block diagram of the Boeing Phased Array Data Acquisition System in given in Figure 77.

The Kulites were connected to Boeing Dual Purpose Power Supplies located in weatherproof 

enclosures near the test stand (Figure 78). Coaxial cables led from the test stand to the host data 

acquisition system, DDS3, which was installed in the control room in the blockhouse at the C-11 test 

stand.
NASA/CR—2014-218088 66

Figure 78.—Weatherproof enclosures for the Boeing dual purpose power supplies.

Figure 79.—Omnidirectional speaker used for array calibration.

4.3.3 Phase Calibration

Speaker calibration was performed with an omnidirectional speaker (Figure 79 and Figure 80) and a 

pipe speaker. As shown in Figure 80 the pipe speaker that was used for high frequency calibration was 

positioned at the opening of the ICD, rather than the center, to optimize the uniformity of its radiation 

pattern over the array.

The omnidirectional speaker was applied for frequencies below 2000 Hz and the pipe speaker was 

used for the frequency range of 2 to 6 kHz. The cross-spectral matrices were calibrated using factors 

determined by ratio of the expected and the measured speaker data.
NASA/CR—2014-218088 67

Figure 80.—Method for positioning the omnidirectional calibration speaker 

in the center of the ICD.

4.3.4 Processing Methods

Two cross spectral matrices were computed for each condition: an “A” matrix with a 10 Hz 

processing bandwidth and 200 center frequencies from 10 to 2010 Hz, and a “B” matrix with a 30 Hz 

processing bandwidth and 200 center frequencies from 30 to 6030 Hz. 

4.3.4.1 Travel Times For Conventional Beamforming

Conventional phased array beamforming (Ref. 1) was used to make images of the noise source 

distributions for selected frequencies and conditions. Flow effects were taken into account by starting 

with computed flowfields for the production and scarf inlets. A beamforming grid in the fan plane was 

defined, and a straight ray between each beamforming grid point and each microphone was computed.

The acoustic travel time for each microphone-grid point pair was approximated by assuming that the 

sound travels along the ray with speed vray c n v

 = + ˆ • , where c is the speed of sound in quiescent air, v

 is 

the local flow velocity, and n

 is the direction of the ray. This approximation was adopted after the more 

rigorous approach of computing the actual, curved, rays proved unworkable in this case. It is believed that 

the curved rays failed to converge because the extreme flow gradients near the inlet lip scattered the rays 

in all directions. The travel times, computed by the straight-ray method (using a code called ray_grid.f), 

were used the compute the phase for beamforming.

4.3.4.2 Spinning Mode Decomposition

To estimate the spinning mode levels, an annular grid was defined in the inlet (Figure 81 and Figure 

82) and used as a source surface. For each spinning order, m, eight radial orders were considered. Source 

modes of the form

















−

− φ = φ

1 0

0 ( , ; , ) cos r r

r r S n m r e n im

Were postulated for n = 0, 1, ..., 8. The cosine factor is intended to account for the radial mode 

variability. The fact that the cosine function differs from the actual mode shape (expressed in terms of 

Bessel functions for some ideal cases) is not believed to be important because the results are ultimately 

summed over the radial modes. For each spinning order, m, the power of the spinning mode was 

estimated by considering each radial order, n, in turn. Given a pair (n,m), the source function S was 

computed at each point in the annular grid. A Rayleigh integral for each microphone location was 
NASA/CR—2014-218088 68

Figure 81.—Position of the annular grid used for spinning mode breakdown.

Figure 82.—Coordinates used for the annular grid source region.

approximated by summing over the grid points and using a free-space, no-flow model for the travel times. 

These integrals became elements of array steering vectors, which were used with the cross-spectral data 

matrices in a beamforming expression. The (spinning mode) beamforming results were summed over n to 

give the required estimates of the radiated spinning mode power.

4.3.5 Test Configurations

Compressor foil counts of the test engine are given in Table 6. These values are used to compute the 

expected spinning mode orders.

The eleven configurations for phased array data are listed in Table 7. The non-scarf inlet is a 

production intake configuration with a flight lip. The scarf inlet is a 13° biplanar design that also has a 

flight lip. Three forward fan cases and several inlet and fan case tape configurations were tested. All 

phased array data set were taken with the inlet keel rotated toward the polar arc of far filed microphones.

r

φ

r0

r1
NASA/CR—2014-218088 69

TABLE 6.—COMPRESSOR BLADE AND VANE COUNTS OF THE TEST ENGINE

Fan Low Pressure Compressor High Pressure Compressor

Blades Vanes Blades Vanes Blades Vanes

Inlet vanes 0 92 38/36

Stage 1 22 60 53 84 40 30/31

Stage 2 91 100 51 26/27

Stage 3 93 104 56 32/33

Stage 4 100 120 53 114

Stage 5 108 110 72 108

Stage 6 94 104 74 120

Stage 7 81 88 78 128

Stage 8 84 112

Stage 9 92 116

Stage 10 90 124

Stage 11 80 110

TABLE 7.—CONFIGURATIONS FOR PHASED ARRAY DATA REPORTED HERE. 

Cfg. Run Corr. rpm

N1C

Obs. rpm

N1 Inlet Forward

Fan Case Acoustic lining and HW tape

1

A07809

A07826

A07830

1802

2450

2900

1828

2494

2967

Prod. Prod. Harwall inlet and F/C

2

A07908

A07918

A07923

1806

2454

2904

1835

2489

2946

Prod. Prod. Production lining

3

A08107

A08217

A08222

1804

2454

2904

1843

2506

2964

Scarf Sim. BOM Hardwall inlet and F/C

4 A08508

A08518

1805

2451

1847

2505 Scarf Sim. BOM Full lining

5

A08608

A08618

A08623

1803

2455

2900

1834

2497

2949

Scarf 2 in. R1 Full lining

6

A08708

A08718

A08741

1802

2450

2907

184

2502

2969

Scarf 2 in. R2 Full lining

7 A08809

A08819

1803

2452

1834

2493 Scarf 2 in. R2 8 3 in. tape strips on the F/C

8 A08909

A08919

1804

2453

1840

2504 Scarf 2 in. R2 3 3 in. tape strips on the F/C

9 A09009

A09019

1804

2453

1849

2513 Scarf 2 in. R2 3 1.25 in. tape strips on the F/C

12

A08309

A08319

A08324

1804

2451

2904

1847

2508

2973

Scarf 2 in. R2 Hardwall lip

14

A08409

A08419

A08419

1802

2452

2907

1849

2503

2969

Scarf Taped HW Hardwall F/C

4.3.5.1 Expected Spinning Modes

Figure 83 and Figure 84 give the inlet spinning modes expected from the engine at approach and 

cutback power according to Tyler-Sofrin theory (Ref. 3). Cutoff effects are not considered in the figure.

For fan modes, it was assumed that three splices are present in the fan case acoustic lining. 

The green diagonal line at the bottom of Figure 83 represents difference tones for various stages of 

the LPC. The blue diagonal line near the bottom is the corresponding feature for the high pressure 

compressor. The group of green diagonal lines cutting across the center of the figure is composed of firstorder LPC tones. The horizontal red lines are fan tones. 
NASA/CR—2014-218088 70

Figure 84 shows the general form of buzzsaw noise. The diagonal red line passing through (0,0) is the 

rotor-locked field for various numbers of blades. The parallel diagonal lines are interactions with the 

hypothesized splices in the lining.

4.3.6 Results

4.3.6.1 Spinning Modes: Approach Power

Figure 85 gives the measured spinning modes for the 11 inlet configurations at approach power. The 

non-scarf inlet in the hardwall configuration (Figure 85(a)) shows the expected co-rotating broadband 

noise near the cutoff line. This exists in the same area of the plot as the LPC difference tones. The first 

two fan harmonics are quite evident. The first order LPC tones are also very clear. There is a suggestion 

of the HPC difference tones.

The treated non-scarf inlet (Figure 85(b)) shows a substantial reduction of the fan noise, but little 

change in the LPC noise relative to the hardwall case.

The hardwall scarf inlet (Figure 85(c)) appears to have more fan tone noise and less fan broadband 

noise than the hardwall baseline. The reduction in broadband fan noise can be interpreted as the scarf 

shielding effect, since the ICD array microphones were clustered on the keel side of the inlet. The LPC 

noise is similar to the previous cases.

The treated scarf inlet (Figure 85(c) to (k)) have dramatically less high frequency LPC noise than the 

non-scarf and the hardwall scarf inlets. The exception is the low-order modes around 2500 to 3000 Hz, 

which are not significantly attenuated.

The fan case changes (Figure 85(c) to (i) and (k)) alter the fan noise slightly. The hardwall lip 

(Figure 85(j)) has increased fan broadband noise relative to the treated lip configurations. 

Figure 83.—Expected spinning modes at approach 

power. (Assumes three splices in the fan case lining.)

Figure 84.—expected spinning modes at cutback power. 

(Assumes three splices in the fan case lining.)
NASA/CR—2014-218088 71

(a) (b)

(c) (d)

Figure 85.—Measured spinning modes for the 11 inlet configurations at approach power.
NASA/CR—2014-218088 72

(e) (f)

(g) (h)

Figure 85.—Continued.
NASA/CR—2014-218088 73

(i) (j)

(k)

Figure 85.—Concluded.
NASA/CR—2014-218088 74

4.3.6.2 Spinning Modes: Cutback Power

The spinning mode plots for cutback power are given in Figure 86. The hardwall non-scarf plot 

(Figure 86(a)) has the expected buzzsaw modes. The production lining gives a substantial, but not 

complete reduction in buzzsaw noise (Figure 86(b)). The hardwall scarf inlet (Figure 86(c)) has a little 

more buzzsaw noise than the hardwall baseline. The treated scarf inlets generally do not radiate buzzsaw 

noise. The exceptions are the configurations with fan case scattering strips (Figure 86(g) to (i)) and 

especially the hardwall lip (Figure 86(j)). Surprisingly, the buzzsaw noise with the hardwall lip appears in 

low order modes. The hardwall fan case (Figure 86(k)) produces a small increase in buzzsaw noise.

Figure 86(j) has increase fan broadband noise relative to the treated lip configurations. 

(a) (b)

Figure 86.—Measured spinning modes for the 11 inlet configurations at cutback power.
NASA/CR—2014-218088 75

(c) (d)

(e) (f)

Figure 86.—Continued.
NASA/CR—2014-218088 76

(g) (h)

(i) (j)

Figure 86.—Continued.
NASA/CR—2014-218088 77

(k)

Figure 86.—Concluded.

4.3.6.2.2 Spinning Modes: Takeoff Power

The spinning mode plots for takeoff power are given in Figure 87. The treated and hardwall plots are 

in the non-scarf case (Figure 87(a) and (b)) are very similar. This indicates that the production lining is 

ineffective at takeoff power. The hardwall scarf inlet has significantly more buzzsaw noise at takeoff 

power that the non-scarf inlet (Figure 87(c)). The treated scarf inlets, except for the one with the hardwall 

lip (Figure 87(d) to (f)) have similar noise characteristics as the non-scarf cases. Figure 87(f) suggests that 

the lip lining is the most important section of lining at takeoff power.
NASA/CR—2014-218088 78

(a) (b)

(c) (d)

Figure 87.—Measured spinning modes at takeoff power.
NASA/CR—2014-218088 79

(e) (f)

Figure 87.—Concluded

4.3.6.3 LPC and Fan Harmonic Noise

Figure 88 to Figure 91 demonstrate that the ICD array finds the LPC spinning modes and higher 

frequency fan harmonics at the expected spinning orders.

Figure 88 shows identifiable spinning modes from as far into the LPC as the fourth stage.

Figure 89 gives conventional beamforming plots at the fan face for modes that were identified as corotating and counter-rotation. The co-rotating modes appear to come from the lower portion of the LPC 

due to the direction of fan and mode rotation (counterclockwise looking into the inlet), and the apparent 

perspective point of the array at the main cluster of microphones on the left side of the inlet. The counterrotating modes come from the top of the LPC.

Figure 90 and Figure 91 show interactions between the fan and the LPC inlet guide vanes, and the fan 

and the first stage LPC rotor.
NASA/CR—2014-218088 80

Figure 88.—Single-frequency spinning mode plots for selected low 

pressure compressor tones (with the hardwall non-scarf inlet) at 

1802 rpm. The mechanical rpm (see Table 7) is 1828 rev./min. 

Interaction modes are shown for LPC stages 1 to 4.
NASA/CR—2014-218088 81

Figure 89.—Fan-face beamforming plots for a co-rotating LPC mode (m = 7 from the first harmonic 

of the second stage of the LPC interacting with the vanes ahead of it), and a counter-rotating LPC 

mode (m = –7 from the third stage of the LPC). Run A08508 (the scarf with the simulated BOM fan 

case) at 1805 rpm corrected, and 1847 rpm mechanical.

Figure 90.—Single-frequency spinning mode plots for interactions between the 

fan and the LPC. In the upper plot, the fan interacts with the inlet guide vanes 

of the LPC. The lower plot shows an interaction between the fan, the first 

stage LPC rotor, and the LPC inlet guide vanes. The same run as Figure 88 is 

represented.
NASA/CR—2014-218088 82

Figure 91.—A fan-face beamforming plot for the fan-LPC IGV interaction show 

in Figure 90. The mechanical rpm is 1847 (it is the same run as Figure 89). 

This is higher than the 1828 rpm of Figure 90, so the frequency is shifted from 

2691 to 2708 Hz. A counter-rotating mode (m = –4) is shown.

Figure 92.—Spinning modes at blade passage frequency for the hardwall 

and treated non-scarf inlet at approach power.

4.3.6.4 Blade Passage Frequency Noise

Figure 92 to Figure 98 give results for the blade passage frequency noise of the various configurations at 

approach power. The hardwall non-scarf inlet has a strong blade passage mode in m = 17-18. (Figure 92). In 

the treated case, the primary mode appears to be m = 9.

The hardwall scarf inlet has elevated BPF energy in a number of modes, including m = 17 (Figure 93).

The treated case is dominated by m = 5.

Beamforming images of the hardwall inlets (Figure 94) show most of the noise coming from the 

general area of the crown in the case of scarf inlet. The noise source distribution appears to have periodic 

structure. The repetitive structure is also evident in the case of the treated scarf inlet and perhaps also the 

treated non-scarf inlet (Figure 95).

Figure 96 to Figure 98 explore the effect of fan case changes on blade passage frequency tones. The 3

and 1.25 in. tape strips increase the noise appreciable (Figure 98 and Figure 99); the other changes do not.

It is unclear if the modes introduced by the tape strips are in the expected orders (22-3k and 22-8k for 

three and eight strips, respectively, and integer k).
NASA/CR—2014-218088 83

Figure 93.—Spinning modes at blade passage frequency for the hardwall and 

treated scarf inlet at approach. 

Figure 94.—Beamforming plots for the hardwall non-scarf and scarf inlets at blade passage 

frequency at approach power.

Figure 95.—Beamforming plots for the treated non-scarf and scarf inlets at blade passage 

frequency at approach power. 
NASA/CR—2014-218088 84

Figure 96.—Spinning modes at blade passage frequency for four fan cases at 

approach power.
NASA/CR—2014-218088 85

Figure 97.—Spinning modes at blade passage frequency for three arrangement of 

scattering strips on the fan case. The R2 case with now strips is shown again 

for comparison.

Figure 98.—Fan face beamforming plots at blade passage frequency for two configurations of 

mode scattering strips.
NASA/CR—2014-218088 86

4.3.7 Conclusions 

The ICD phased array can measure low-pressure compressor noise in the expected spinning modes.

The results for blade passage frequency are promising, but not completely satisfactory. This is probably 

due to the scattering of the modes by the nonuniform inlet and flow geometry. At approach power, the 

inlet noise of the PW4098 has a strong contribution from the low pressure compressor. Blade passage 

tones and broadband fan noise are also important. High pressure compressor noise may play a role. Blade

passage frequency noise for the hardwall baseline inlet is dominated by the mode m = 17-18. There is no 

obvious mode in the treated case. The blade passage frequency noise of the treated scarf inlet is strongly 

concentrated around m = 5. Beamforming images show a periodic source structure that may be consistent 

with generation of this mode. Inlet lip lining is very effective in attenuating buzzsaw noise at low 

spinning orders. The production lining is ineffective at takeoff power. The production lining is largely 

ineffective for low-pressure compressor noise at approach power.

4.4 PW4098 Blade-Mounted Transducer and Mode Array Tone Diagnostic Test Results

4.4.1 Data Acquisition and Instrumentation

4.4.1.1 Data Acquisition System

Figure 99 shows a schematic of the data acquisition system. This system was comprised of a network 

of two laptop computers (233 MHz) and one desktop computer (133 MHz) with an external Iomega Jaz 

drive. Connection to the Boeing network was accomplished by modem. 

Figure 99.—Schematic of Instrumentation setup.
NASA/CR—2014-218088 87

A Sony SIR-1000 system with a 32-channel capacity was used for data acquisition. The system was 

capable of measuring 32 channels simultaneously at a sampling rate of 48 kHz. Data was digitally 

recorded onto a Sony 8 mm tape, which was played back through an AIT (Advanced Intelligent Tape) 

drive unit, (model number SDX-S300C) and converted to an ASCII file on the PC.

4.4.1.2 Blade-Mounted Transducers (BMT)

Eight Kulite pressure sensors (Model LQ-47-50A) were mounted to the pressure side of two fan 

blades at the locations shown in Table 8. A nominal distance of 0.1 in. was assumed to be the gap 

between the blade tip and the wall. The distance from the blade tip to the transducers was therefore the 

value shown in Table 8 are -0.1 in. Four transducers were mounted on the number-1 and number-22 

blades as shown in Figure 100. The telemetry system only allowed six signals to be transmitted at any 

time, but due to some complications, a maximum of five BMT’s were used. The increased boundary layer 

thickness in the scarf inlet required a distribution of Kulites extending to greater distance from the wall 

than the Production inlet.

Before the transducers were bonded to the blades, the blades were heated in an autoclave to remove 

any dirt and oil that may have contaminated the blade surface. The leads from the sensors were routed as 

shown in Figure 100 along the blade mid-chord, and into the hub. 

The Pratt & Whitney telemetry system was mounted in the spinner in order to transmit the pressure 

signals to an antenna in the inlet. The antenna was secured to the inlet with Rhino tape in the keel region 

as shown in Figure 101.

TABLE 8.—BMT LOCATIONS ON FAN BLADES

Blade 1 Blade 22

Transducer Distance from wall,

in.

Transducer Distance from wall, 

in.

1A 0.5 2A 1.0

1B 1.5 2B 2.0

1C 2.5 2C 4.0

1D 3.25 2D 8.0

Figure 100.—Locations of BMT’s on Blades 1 and 22 (not to scale).
NASA/CR—2014-218088 88

Figure 101.—Photograph of scarf inlet indicating the location of the telemetry antenna.

4.4.1.3 Inlet Mode Array

In the second half of testing, May to June 1999, an adapter ring was made to allow the production 

case to be used with the scarf inlet. This provided an opportunity to add a circumferential microphone 

array in the inlet for measuring the acoustic modes. A 26-transducer, half-circular array spanning the top 

half of the engine was designed into the adapter ring, which can be seen in Figure 102.

The transducers were located at the attach flange (STA 121) which is approximately 13 in. from the 

fan face. The transducers are evenly spaced 6.92° apart spanning half the circumference of the inlet.

Special pressure transducers were used, which had very little phase variations between them. These 

transducers were developed to support other NASA sponsored work to measure broadband modes for the 

Source Diagnostic Test currently scheduled for February of 2000. The transducers were a slight variation 

of Kulite Model XCQ-125B-093-15A having an RTV layer over the diaphragm instead of a protective 

screen. The piezo-resistive transducers are capable measuring the pressure without the need for a phase 

calibration between the sensors. 
NASA/CR—2014-218088 89

Figure 102.—Location of the mode transducers.

Figure 103.—Ideal Inlet Mode Data System block diagram.
NASA/CR—2014-218088 90

A disadvantage of these sensors is that they are very easily damaged because the diaphragm is 

exposed. Approximately 15 transducers were lost during testing. Most were lost due to a design flaw in 

the adapter ring that didn’t allow for trouble-free sensor installation.

The sensors were wired to the bridge/balance power supplies located in a room near the test stand (see 

Figure 102) using 126 ft of wire. The voltage signals from the power supplies were then connected to the 

Control Room over the P&W existing ground lines by using specialty cables. A Sony SIR-1000 recorder 

was used to simultaneously record 8-sec samples at 48 kHz for the 31 input channels onto an 8 mm 

Advanced Intelligent Tape (AIT). 

A Sony AIT Streamer Drive was used to transfer the data from tape onto a network of personal 

computers for processing. Analysis codes were written at Boeing to process the data to produce pulseand non-pulse synchronous data.

4.4.1.4 Data System Trigger

A “short tooth” gear was used as the trigger signal for data acquisition and pulse-synchronized 

averaging. Figure 104 is a typical trace of the trigger signal as measured by the data acquisition system. 

Table 9 provides the details of the pulse signal.

Figure 104.—Typical trace of trigger signal. (a) 587 rpm, Volts versus time, (b) 1833 rpm, Volts 

versus time, (c) 587 rpm, Volts versus angle, and (d) 1833 rpm, Volts versus Angle.

TABLE 9.—DETAILS OF TRIGGER SIGNAL

587 rpm 1833 rpm

Rise Fall Rise Fall

Edge Slope, V/ms 113.6 -122.8 108.5 -111.0

Edge Width, ms 0.0417 ms 0.0417 ms 0.0417 ms 0.0417 ms

Edge Width, deg. 0.147° 0.147° 0.46° 0.46°

Pulse Width, sec 0.792 ms 0.604 ms

Pulse Width, deg. 2.79° 6.64°
NASA/CR—2014-218088 91

4.4.2 BMT Data Analysis

Voltage signals from the pressure transducers and trigger channel were recorded simultaneously. 

Once the data were recorded, they were then re-sampled at the desired angular resolution, which allowed 

pulse-synchronized mean and rms pressures to be calculated. By knowing the location of the instrumented 

blade, these values could then be converted from their respective time traces to their physical position 

relative to the inlet. A complete data set covered 8 sec of run-time for a single engine operating-point, 

which amounted to approximately 213 revolutions at 1600 rpm, and 386 revolutions at 2800 rpm.

A trigger level of 2 V on the falling edge was used as the starting point for each block of data, which 

typically consisted of one revolution. Since the operating speed of the fan is not perfect, small variations 

in fan rpm occur for each revolution. In order to ensure that each revolution in the averaging process had 

data points at exactly the same angular location, the time between the closest acquisition sample and the 

2V location was calculated for each rotation. From this the entire time series for that revolution was 

shifted and re-sampled with the desired angular resolution. Since the original data set was sampled at 

48 kHz, the worst angular resolution possible was at the highest rpm; e.g., at 2800 rpm, this corresponded 

to 1028 pt/rev, or an angular resolution of 0.35°. The majority of the re-sampling was performed with a 

resolution of 0.5° or 1°.

Four fundamental analyses were performed on the data as shown in Figure 105, Pulse-synchronized 

mean pressures, pulse-synchronized rms pressures, power spectrum of the pulse-synchronized mean 

pressure, and the mean power spectrum of the individual spectra from each rotation. 

The pulse-synchronized mean pressure was calculated by averaging the pressure at a given blade 

location over the entire number of revolutions in the data set. Since the transducers measured the unsteady 

pressure (AC coupled), the time average around the circumference is theoretically zero. This analysis 

provides a measure of the deviation of the pressure at a specific angular location relative to this zeromean pressure for one revolution of the fan. The absolute pressure sensed by the blade is therefore equal 

to the mean pressure in the inlet plus this calculated mean pressure.

The RMS pressure at a specific angular location was then calculated by using the previously 

calculated mean pressure at a specific location and the pressures from each revolution for that location. 

The RMS pressure provides an indication of the unsteady pressure at each angular location relative to the 

local mean pressure at that angular location.

The power spectrum of the mean pressure was calculated by applying a temporal Fourier transform to 

the mean pressure signal. The frequency domain was then converted to distortion orders around the inlet. 

A distortion order of 1 implies that the BMT senses a one-cycle distortion as the fan makes one 

revolution. Since the time series was averaged prior to the Fourier transform, randomness of the time 

series was removed and hence lower “broadband” levels occur for this analysis. 

By applying the Fourier transform to a block of data, which corresponds to one fan revolution, the 

different distortion modes from each revolution are retained. When these spectra are averaged, a higher 

broadband spectrum will result in significant differences exist between the individual fan rotations.
NASA/CR—2014-218088 92

Figure 105.—Schematic of BMT data analysis.
NASA/CR—2014-218088 93

4.4.2.1 BMT Results

The mean pressure profiles for the range of BMT locations is shown in Figure 106 for the 1620 N1C 

condition for the production inlet with and without the ICD, and the scarf inlet with the ICD. (The scarf 

was never operated without the ICD.) These plots show the mean pressure around the circumference of 

the inlet for a particular radial location. The value of the pressure indicates the variation of pressure from 

a mean value of zero around the circumference of the inlet. Each trace shows the characteristic once-perrevolution variation in the static pressure, which has been associated with the inlet droop. The amplitude 

of this variation increases with fan speed.

The production inlet without the ICD shows the lowest variations in mean pressure around the 

circumference of the inlet for the BMT’s within 1 in. of the inlet wall. When the ICD is installed on the 

production inlet, large stationary disturbances occur in the boundary layer. The scarf inlet shows some 

regions of stationary disturbances that are on the order of those seen in the production inlet when the ICD 

is installed, with the dominant feature being the large distortion in the crown region. This distortion exists 

for all BMT locations out to 8 in. from the inlet wall.

The unsteady RMS pressure traces are shown in Figure 107 for the 1620 N1C condition. These values 

indicate the amount of variation in the static pressure at a given radial and angular position around the 

inlet. When the ICD is installed, the unsteady pressures decrease at all circumferential and radial locations 

for the production inlet. The scarf inlet shows the highest unsteady pressures in the crown region, which 

will later be shown to be due to the lateral motion of the boundary layer flow. Once out of the boundary 

layer, both inlets show similar levels of unsteady pressure, although the scarf still has some evidence of 

increased unsteadiness near the crown.

Figure 106.—Mean pressure traces for BMT’s at 0.5, 1.0, 2.0, and 8.0 in. for the 1620 N1C condition.
NASA/CR—2014-218088 94

Figure 107.—RMS pressure traces for BMT’s at 0.5, 1.0, 2.0, and 8.0 in. for the 1620 N1C condition.

The once-pre-rev variation in the mean pressure around the inlet was removed by performing a 

Fourier transform of the mean pressure, setting the fundamental mode to zero, and then taking the inverse 

Fourier transform. These data are shown in Figure 108 in the form of contour plots of the mean pressure 

and the associated unsteady pressure for the 1620 N1C condition. The plots have been stretched radially 

to accentuate the contours, and thus are not to scale. The no-ICD case only had BMT’s out to 2 in. from 

the wall, while the other cases had BMT’s out to 8 in. from the wall. Without the ICD, the unsteady flow 

is the dominant feature in the keel region of the inlet. When the ICD is installed on the production inlet, 

some steady distortions are seen in the inlet boundary layer, while the unsteadiness is greatly reduced. For 

the scarf inlet the large distortion near the crown dominates the steady pressure, and shows a region of 

unsteadiness at ±90° around the crown. The steady distortions in the keel region that were seen in the 

production inlet with the ICD are not apparent with the scarf. Since the scarf inlet has a flight lip, the 

boundary layer develops over a region with a much larger pressure gradient than the bellmouth on the 

production inlet. Because of this, the scarf boundary layer is more susceptible to disturbances that would 

amplify the boundary layer instabilities, and hence cause a more rapid growth of the boundary layer, 

possibly smearing out the steady distortions seen in the production inlet.
NASA/CR—2014-218088 95

Figure 108.—Mean and unsteady pressure contours for the 1620 N1C condition.

Waterfall plots of the measured pressure are shown in Figure 109 for the first 50 revolutions of the fan 

for BMT 1A, which was 0.5 in. from the inlet wall. The No-ICD case shows structures that exist over 

many revolutions of the fan, yet shift in the azimuthal direction during operation. From 0° to 30° for the 

first 20 revolutions a structure is seen which moves in a counterclockwise direction. Near the 26th to 36th

revolutions in this same angular region, another structure moves in the opposite direction. Similar 

structures are seen over the full circumference of the inlet when the ICD is not installed. The movement 

of these structures is what creates the high unsteady pressures that were measured at a given angle by the 

BMT’s. Since these structures exist over several rotations of the fan, their interaction with the fan blades 

will result in BPF noise. The movement of these disturbances will modulate the tone spreading the BPF 

energy to the neighboring sidebands.
NASA/CR—2014-218088 96

Figure 109.—Waterfall plots of raw pressure signal for BMT 1A (0.5 in. from wall) at 1620 N1C; 

(a) production inlet, no ICD, (b) production inlet with ICD, and (c) scarf inlet with ICD.

When the ICD is installed on the production inlet (Figure 109(b)), the amplitude and unsteadiness of 

these coherent structures at a given angle diminishes. The only remaining disturbances are quite steady in 

space and exist most strongly in the keel region. These structures are coherent over numerous revolutions 

and will generate BPF tones as in the non-ICD case, but since they are stationary, the resulting tone will 

be narrower.
NASA/CR—2014-218088 97

When the scarf is installed with the ICD as seen in Figure 109(c) high-amplitude steady distortions 

are seen in the crown region. On either side of the crown coherent structures exist which are drawn to the 

crown region because of the circumferential pressure gradient. The lateral velocity of these distortions 

was calculated to be around 8 ft/s, or 4 in. of lateral movement for each fan revolution. These structures 

and the dominant disturbance at the crown interact with the fan blades to generate the BPF tone. The 

modulations in space and amplitude of these disturbances will also broaden the BPF tone spreading 

energy into the sidebands.

By taking the Fourier transform of the pulse-synchronized mean pressure signal, the inlet steady 

distortion modes may be calculated. These are shown in Figure 110 for the 1620 N1C case and describe 

the amplitude of the steady distortion modes that are encountered by the fan blades. These interactions 

will generate spinning modes based on the Tyler-Sofrin interaction: m = σB ±k, where m is the spinning 

mode, σ is the harmonic of the blade passage frequency, B is the number of blades, and k is the inlet 

distortion mode order. At 1620 N1C, spinning modes less than about 14 will propagate, indicating 

distortion orders from 8 to 36 will contribute to the BPF.

For the data shown in Figure 110, the production inlet without the ICD shows the lowest amplitude 

steady distortion modes because most of the disturbances seen by the fan blades are unsteady. When the 

ICD is installed on both the production and scarf inlets, the unsteady atmospheric distortions are broken 

up and dissipated by the turbulence. An attempt was made to relate the spacing of the ICD panels around 

the nacelle to these distortion orders, but no clear connection was found. The steady distortions from the 

scarf inlet show up in the low orders because of the impulse-like structure of the crown distortion whose 

energy is spread among all of the modes.

As the distance from the wall increases, the amplitudes of the higher-order modes decrease. Peaks are 

seen in the modal spectrum around 90 to 110, which are due to the interaction of the fan with the acoustic 

field of the low-pressure compressor blades. These acoustic fields not only interact with the sensors 

farthest from the wall, but also with those that are well within the boundary layer.

Figure 110.—Power spectra of mean spectra for BMT’s at 0.5, 1.0, 2.0, and 8.0 in. for the 1620 N1C 

condition.
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Figure 111.—Averaged power spectra of phase-locked pressures for BMT’s at 0.5, 1.0, 2.0, and 8.0 in. for 

the 1620 N1C condition.

If the power spectrum is calculated for each revolution and then averaged, the slowly moving 

disturbances that contribute to the unsteady pressures will remain in the distortion mode amplitudes. 

Figure 111 shows the amplitudes of the distortion orders for a range of radial locations. For this analysis 

the production inlet without the ICD shows the highest amplitudes since each revolution of the fan 

experiences more abundant and higher-amplitude distortions than when the ICD is installed. This is true 

for all radial locations and fan speeds. The scarf inlet shows moderately high distortion mode amplitudes 

since unsteady distortions exist in the crown region. In addition to this, energy from the impulse-like 

distortion in the crown region of the scarf will also be spread out over all other distortion modes.

4.4.2.2 Rotated Scarf Inlet Results

For the previous data, the inlet crown was at approximately the 10:00 position relative to the engine. 

The scarf inlet was also tested in a rotated configuration on the engine such that the crown was at the 

12:00 position and the keel was closest to the ground. The distortion in the crown region however, did not 

rotate the same amount as the inlet. This is illustrated in Figure 112 for the mean pressure contours with 

the first distortion order removed for the 2450 N1C condition. These plots also indicate a 2nd order 

distortion that does not rotate with inlet and exists for both the production and the scarf inlet. The highpressure regions are near 50° and 230°, while the low-pressure regions are near 140° and 320°.
NASA/CR—2014-218088 99

Figure 112.—Mean pressure contour plots at 2450 N1C, k = 1 distortion mode removed; (a) production inlet with ICD, 

(b) scarf inlet with ICD, and (c) rotated scarf inlet with ICD (crown at 0°).

Figure 113.—RMS pressure contour plots at 2450 N1C; (a) production inlet with ICD, (b) scarf inlet with ICD, and

(c) rotated scarf inlet with ICD (crown at 0°).

The mean and rms pressure profiles for each configuration are shown in Figure 114 and Figure 115

where the data from the rotated scarf has been shifted 84° to coincide with the inlet orientation of the 

other configurations. The mean pressures show that not only is the scarf crown-distortion shifted, but the 

droop induced once-per-rev variation has been altered as well. The unsteady pressures also show a 

different character for the rotated condition.

The second order distortion and the displacement of the crown distortion in the scarf could be an 

effect of the pressures imposed by the secondary flow path in the engine due to the pylon and support 

struts. Other factors contributing to these differences may be the external support structure, ground 

effects, or the aft-fan acoustic barrier.
NASA/CR—2014-218088 100

Figure 114.—Mean pressure comparisons for the rotated scarf inlet at four BMT locations.

Figure 115.—RMS pressure comparisons for the rotated scarf inlet at four BMT locations.

4.4.2.3 Conclusions

The ICD has been shown effective in eliminating the large-scale atmospheric disturbances, and the 

accompanying unsteady pressures they create. The presence of the ICD however, did introduce some 
NASA/CR—2014-218088 101

steady distortions in the boundary layer of the production inlet that may contribute to the BPF tone. These 

steady distortions were not present with the scarf inlet, although a large distortion in the crown region 

existed due to separated flow aft of the throat. The lateral pressure gradient in the boundary layer of the 

scarf inlet caused axially coherent structures to migrate towards the crown which could cause a modulated 

BPF tone and subsequent harmonics which would be broader in frequency than the production inlet tone.

4.4.3 Flow Visualization

Flow visualization was used to indicate flow patterns and regions of separation in the scarf inlet. 

Retroreflective string was glued to the inlet with a hypodermic needle in a 2- by 2-in. grid in the crown 

region from the highlight to the A-flange The tufts were visualized by mounting a miniature camera and 

light source on the ICD and recording the tuft motion on videotape. Figure 116 shows a wide-angle view 

of the inside of the ICD. When the ambient air had sufficient humidity the moisture in the air would 

condense due to the acceleration of the flow. An example of the streaklines of the condensed air can also 

be seen at this fan speed. The video indicates motion of these of these structures, but it is not clear 

whether it is a similar motion of the structures that are seen in the waterfall plots, or merely a 

manifestation of the variation in air density combined with the shutter speed of the camera.

Figure 117 shows a close up of the tuft motion for the 2452 N1C condition. Downstream of the 

throat, tuft motion was seen for all fan speeds indicating separated flow. At the inlet lip the tuft motion 

was sufficient to scrub off the paint and cause breakage of the tufts indicating more separation and 

unsteady flow in lip region as well. No indication of flow direction other than axial was observed. From 

the waterfall plots, the transverse flow speeds of the convected disturbances were considerably less than 

the freestream, which would produce flow angles much less than 1°.

Figure 116.—Digitized image of retroreflective tufts and condensation flow visualization in the 

scarf inlet at 2800 rpm.
NASA/CR—2014-218088 102

Figure 117.—Digitized image of retroreflective tufts in the scarf inlet at 2452 rpm.

4.4.4 Inlet Mode Array Analysis

The data was processed and plotted in several ways to help understand the source of the tone noise. 

Each method showed different aspects of the data, and thus all were useful in the analysis. Additionally, 

combining the results of the mode array with the results from the blade mounted transducers (BMT) 

allowed even greater understanding of the BPF tone generation mechanisms.

Much of the data shown in this document is from the lowest power setting (N1C = 1620 rpm). This 

was a particularly useful power setting because the BPF was cut-off and theoretically, there should not 

have been any BPF noise. The primary purpose of the mode array and BMT transducers was to determine 

the source of the BPF noise at subsonic tip speeds for the production inlet and determine the cause of the 

increase in BPF noise at subsonic tip speeds of the scarf over the production inlet.

4.4.4.1 Mode Array Pressure Traces

Analysis of data started with using the once-per-revolution pulse to re-sample the data at specific 

angular locations of the rotor. This required using a digital linear-phase finite impulse response (FIR) 

filter to first remove frequencies above the new Nyquist frequency in the data. A polynomial splineinterpolation routine was used to re-sample the data. The data was over-sampled by a factor of 

approximately 10 so that interpolation did not lead to significant errors. 

Average pressure traces representing one rotation of the rotor were produced by averaging each resampled pressure trace over a period of eight seconds for each condition. For the low powers there were 

approximately 210 revolutions at the lowest power of N1C = 1620 rpm and approximately 380 

revolutions at the highest power of N1C = 2800 rpm. By performing this pulse-synchronized averaging 

the random part of the signal, (unsteady between fan rotations), was reduced by 10*LOG10(N) dB, where 

N is the number of blocks in the average.

The pulse-synchronized averaging process can dramatically reduce the non-synchronous noise and 

improve the ability to view “pure” tones. In this context, pure tones are tones that are completely 

deterministic and therefore have infinitesimally small bandwidth. 

Figure 118 shows a comparison of the average pressures measured over one rotation for the scarf and 

production inlets as a function of power upstream of the keel. Note the pressure increase as each blade passes 

the transducer, which provides a very clear blade-to-blade pressure variation associated with each blade. At 
NASA/CR—2014-218088 103

the subsonic tip speeds, below N1C = 2200 rpm, the rotor locked field is cut-off and an m = 22 pattern should 

not propagate. However, the mode array is sufficiently close to the fan that the cut-off rotor-locked field 

dominates the pressure traces. In hindsight, it would have been beneficial to have the measurement array 

farther away from the fan to minimize the effect of the decaying, rotor-locked pressure field.

Figure 118.—Production versus scarf pressure profile at the keel.
NASA/CR—2014-218088 104

Figure 119 shows the same comparison at the crown. Notice that the production and scarf pressure 

traces differ significantly in this region at the higher powers. Evidence from the flow visualization and 

BMT data indicate separated flow in the scarf inlet in this region.

Figure 119.—Production versus scarf pressure profiles at the crown.
NASA/CR—2014-218088 105

Figure 120.—Pulse synchronized average spectra, and pulse synchronized spectra. 

4.4.4.2 Mode Array Power Spectra

The power spectra of the pulse-synchronized average pressures produces spectra which have reduced 

amplitudes between the BPF harmonics since the randomness is eliminated in the time-series averaging. 

This type of averaging is defined as, “pulse-synchronized average spectra”. By averaging the spectra from 

each individual rotation, (pulse-synchronized spectra) the randomness from non-rotor-locked sources is 

maintained in the spectra as well as the deterministic sources, and hence higher levels appear between the 

BPF harmonics. The difference between these two processing techniques is evident in Figure 120. The 
NASA/CR—2014-218088 106

BPF (m = 22) tones generate similar levels in both processing techniques, with the difference being a 

1.76 dB decrease in the peak level accompanied by a widening of the tone in the pulse-synchronized 

spectra due to the use of a Hanning window. 

The difference between the two processing methods is particularly apparent in the broadband levels 

where one would expect the spectra to be dominated by non-synchronous sources. There is approximately 

a 17 dB difference between the two processing methods for this condition which corresponds to 

10*LOG10(number of averages). At the lower frequencies, multiple pure tones exist, although these 

should not propagate at this low power condition.

The production inlet has relatively little change in levels from the crown to the keel as shown in 

Figure 121(a), while the scarf inlet has higher levels at the crown than at the keel as seen in Figure 121(b). 

The scarf inlet BPF tone is wider than the production inlet and the disparity increases as the crown is 

approached. This suggests there might be some tone modulation, causing spreading of energy in 

frequency, for the scarf compared to the production inlet.

Figure 121.—Pulse-synchronized spectra at all microphone locations (a) production and (b) scarf.
NASA/CR—2014-218088 107

4.4.4.3 Phase and Amplitude Modulation

Phase and amplitude modulation charts of the BPF tone were created from the FFT of the time-traces 

for each revolution. At a N1C of 2200 rpm the rotor locked pressure field propagated which increased the 

signal to noise ratio for the BPF tone. Figure 122 shows the amplitude and phase modulations of the BPF 

with respect to time for all of the transducers. These values were normalized by their average phase and 

the amplitudes, respectfully.

Figure 122.—Production and scarf modulation data.
NASA/CR—2014-218088 108

The phase modulation of the production inlet shows similar modulation across all the sensors, which 

results from a source that effects all sensors simultaneously. Two possibilities exist. The first would be a 

problem with the trigger, although this is unlikely. The second possibility is that there is a large distortion 

across the inlet, which affects the phase at all transducers, possibly resulting from atmospheric variations. 

A similar phase modulation is observed with the scarf inlet, although it is somewhat masked by the 

differences that exist between the transducers. The production inlet also shows small differences in phase 

and amplitude modulation between transducers, while the scarf inlet shows considerable modulation. The 

largest variations for the scarf inlet occur in the crown region where it has been postulated that this 

variation of the BPF modulation is related to the cause of the increase in BPF noise of the scarf inlet.

4.4.4.4 Spinning Mode Charts

The pressure signals can be used to generate mode charts by applying a discrete Fourier transform 

(DFT) in both frequency and space (or angular position). The resulting spectra separate the pressure 

signals into frequencies and acoustic spinning modes. Figure 123 shows an example of a typical mode 

plot for the production inlet at 1620 N1C. The ordinate shows the frequency, and the abscissa contains the 

spinning orders, (m-orders). Since the array only covered half of the inlet circumference, only the even morders are can be calculated. Higher resolution methods were attempted to give some definition of the odd 

harmonics. These methods, however, generally failed for two reasons: 1) there was a very strong BPF, 

which leaked into the odd harmonics, 2) any spinning mode could exist so there were 2xN variables and 

1xN pieces of data. Higher resolution methods likely would have worked had the sensors been further 

away from the source (to remove cut-off BPF noise), thus limiting the number of existing modes. Several 

horizontal lines are noticeable in Figure 123 representing different BPF and compressor tones and 

harmonics. There is a measured BPF tone at approximately 620 Hz, although it is theoretically cut-off at 

this low operating condition. 

Modes near cut-off propagate down the duct very slowly and give increased pressure amplitudes at 

the wall transducers. This phenomenon generates a “V” in the plot which is originates at the m = 0 mode 

at zero frequency. Cutoff for this particular figure at the BPF is at m = 16. Another artifact of the data is 

the preponderance of co-rotating modes for the broadband noise. This is consistent with earlier work of 

Joppa (Ref. 1).

Figure 123.—Mode plot for the production inlet at N1C = 1620 rpm.
NASA/CR—2014-218088 109

Figure 124.—BPF mode plot for the production inlet at N1C = 1620 rpm.

Figure 124 represents the mode amplitudes from the horizontal line at the BPF of Figure 123. Notice 

that the levels are very high for the cut-off m = 22 spinning order. Due to the close proximity of the array 

to the fan (about 13 in.), there is still significant energy present in this cut-off mode. Clearly, future mode 

measurements should be made upstream of the A-flange so that cut-off mode energy has sufficiently 

attenuated before the measurement transducers.

Figure 125(a) and (b) show the mode charts for the production and scarf inlets at 1620 N1C, 

respectively. Note that the levels are the same except near the negative spinning orders at the 620 Hz 

BPF, where the scarf inlet shows additional noise. Careful examination of this source shows that it seems 

to spread out in frequency as the spinning orders move away from the m = 22 spinning order. This is an 

indication that the additional noise source of the scarf inlet may be due to an interaction of the rotor with 

an unsteady disturbance with a long axial length scale. Figure 126 shows the same comparison at 2450 

N1C with a BPF near 920 Hz. Note that the extra source is still visible although the far-field data showed 

little or no increase in BPF noise at this frequency. Since the rotor-locked pressure field cuts-on at this 

speed, the dominant m = 22 can propagate, and hence dominates the far-field levels. Therefore, this 

source is only important at subsonic tip speeds (below N1C of 2200 rpm).

Pulse-synchronized averaging can also be used for mode data to produce spectra that remove the 

random part of the signal as was done with the spectra from the individual sensors. Figure 127 shows a 

comparison of the BPF for the scarf and production inlets using the two spectral averaging methods at 

1620 N1C. The production inlet shows the pulse-synchronized average spectra to be similar to the pulsesynchronized spectra. This suggests that the source of BPF noise for the production inlet is due to a 

deterministic mechanism.

A comparison of the production and scarf pulse-synchronized average spectra show similar levels at 

the cut-on spinning orders. This suggests the steady-state deterministic sources for these two inlets are not 

greatly different. This is contrary to expectations looking at the averaged pressure profiles for the BMT 

data. It suggests that the steady modulations in the boundary layer do not appreciably affect the acoustic 

pressures measured at the mode transducers.
NASA/CR—2014-218088 110

Figure 125.—Mode plots at N1C 1620, (a) production and (b) scarf.
NASA/CR—2014-218088 111

Figure 126.—Mode plots at N1C 2450, (a) production and (b) scarf.

A comparison of the two averaging methods for the scarf inlet spinning mode amplitudes show that 

there is an additional source, which is non-deterministic in nature that is particularly apparent at the 

negative spinning orders. This suggests that the additional scarf inlet noise mechanism does not exist in 

the production inlet. 

The BMT data shown in Figure 109(c) clearly confirms the existence of flow disturbances with long 

axial length scales. This explains the source of the additional BPF noise of the scarf at subsonic tip 

speeds. Note that this type of source is only significant at subsonic tip speeds. When the tip speeds are 

sonic and the rotor locked field propagates, the rotor locked field, m = 22, is much stronger than the other 

spinning orders and controls the far-field levels. The far-field data clearly show an increase in noise of the 

scarf over the production until the rotor-locked field is cut-on. After which, the levels are similar.
NASA/CR—2014-218088 112

Figure 127.—Comparison of pulse and non-pulse synchronized data.

4.4.5 Conclusions and Recommendations

The mode array, in combination with the BMT’s and the flow visualization techniques, provided 

insight into some of the BPF noise-generation mechanisms for the scarf and production inlets. These data 

show the additional BPF noise for the scarf over the production inlet at subsonic speeds is due to a flow 

problem at the crown. The current scarf inlet, without a bellmouth, separates at the crown, which causes 

vortices to be shed into the fan. These vortices have long axial length scales, which are not stationary and 

hence, generate a non-deterministic BPF noise source in spinning orders that propagate out the nacelle. 

High peak Mach numbers around the inlet lip of the scarf inlet can be reduced with the use of a 

bellmouth, thus reducing the possibility of separated flow. Recent CFD calculations indicate that the 

separated flow in the scarf inlet will be eliminated with forward flight. Additional work is underway to 

determine the possibility of testing the scarf inlet with a bellmouth inlet.

The production inlet with the ICD did not show evidence of a non-deterministic BPF noise source in 

the mode array or BMT data. Since the production inlet uses a bellmouth inlet which reduces the Mach 

number near the lip, the probability of flow separation is minimized thereby eliminating the existence of 

these unsteady flow distortions. In addition, CFD results have shown the scarf inlet to have higher peak 

Mach numbers than the production inlet due its flow lines. Recent CFD runs have shown some evidence 

that the scarf inlet would not have the separation BPF noise source if it was in forward flight or had a 

bellmouth inlet. 

The cause of the BPF noise for the production inlet at subsonic speeds was never fully determined. 

The data was able to show that the production inlet BPF noise is due to a deterministic noise mechanism, 

but the source of the mechanism is not conclusive. The BMT data show steady distortions which exist in 

the boundary layer, yet it is not clear whether the noise generated from the interaction of the fan with 

these disturbances is sufficient to generate the observed BPF tones. The mode array data, however, shows 

that the noise source is spread out across all the propagating spinning orders. This implies that the source 

is not comprised of one interaction that can be determined and fixed. The source seems to be comprised 

of many propagating spinning orders. Mode measurements and/or BMT data during 

acceleration/deceleration and with the inlet rotated may provide the additional data to needed to help 

understand this noise source. These types of data were not taken during this test. 
NASA/CR—2014-218088 113

The mode array data did not provide any evidence of large differences between the production and 

scarf deterministic noise sources. However, the BMT data clearly showed large steady distortions in the 

pressures measured on the blades for both inlets. There is not currently enough data to determine the 

amount of additional noise due to the steady distortion as measured with the BMT’s. The data clearly 

show that the unsteady part of the noise dominates the differences. Possibly, some of the fan-noise 

prediction codes could be used to predict the relative levels of the steady and unsteady noise sources from 

the BMT data to compare these sources. The data suggests that the steady source increase is insignificant 

in comparison. 

Clearly the mode array, BMT’s, and the visualization techniques are useful in helping understand 

complex BPF noise mechanisms. It is unlikely that the source of the BPF noise for the scarf inlet would 

have been discovered and so fully understood without these techniques. Likewise, there now exists a 

greater understanding of the type of source that needs to be eliminated in order to reduce subsonic BPF 

noise of current engines.

5.0 Conclusions

For convenience the test conclusions listed in various portions Section 4 are summarized here:

5.1 Far Field Data

• The treated scarf inlet with the advanced (R1) FFCC is 2 to 4 EPNdB quieter in inlet fan noise 

than the production inlet and fan case at approach and cutback powers. Approximately 80 percent

dB nacelle suppression improvement was observed at approach power and 50 percent dB 

suppression improvement at cutback power.

• The Scarf inlet with advanced FFCC has a significant BPF distortion tone, preventing larger 

benefits. Without the extra BPF noise an additional 1 EPNdB in noise reduction was expected.

• The hardwalled scarf inlet has higher buzzsaw noise levels than the production inlet but treatment 

suppressed it very well.

• The bottom 90° of diffuser liner is ineffective at all conditions.

• The scarf lip liner effectiveness at cutback power is about the same per unit length as the diffuser 

treatment.

• The scarf lip liner at approach power is effective over a wide frequency range, but not for the 

BPF distortion tone thus preventing a significant EPNdB benefit.

• At approach power both advanced FFCC liners (R1 and R2) have the same effectiveness per unit 

length as the scarf inlet liner.

• At cutback power the higher resistance fan case liner (R1) showed a significant benefit while the 

lower resistance liner (R2) was detrimental, primarily due to impedance non-uniformities.

• Treatment uniformity near the fan is very important. Simulated production type hardwall splices 

in the FFCC liner show large penalties of up to 4 EPNdB at cutback power for eight 3 in. splices. 

Penalties increase with both splice number and width. The mechanism is scattering, by 

impedance discontinuities, of high level but fast decaying blade pressure fields at BPF (m = B) 

into high level propagating modes.

• The treatment uniformity further forward in the inlet is not critical. Penalties associated with axial 

or circumferential splices are associated only with lost treated area.

• There is no apparent benefit of the scarf keel being oriented toward the microphones, relative to 

being oriented 90° away. This may be due to the effect of the scarf geometry on distortion tone 

generation or on azimuthal buzzsaw noise variations.

• Aft fan noise is not influenced appreciably by the scarf inlet. 
NASA/CR—2014-218088 114

5.2 RDIFF Prediction Comparisons 

At the approach power condition it is concluded that the RDIFF predictions for tone and broadband 

lining attenuation generally were within about 15 percent of the measured data in the direct radiation field 

but only within about 50 percent in diffraction dominated field (aft of 70°). However, the very high 

frequency lining attenuation (8 to 10 kHz) was not accurately predicted. The scarf shape shielding effects 

were predicted well except for in the diffraction-dominated field also. The RDIFF prediction process of 

applying the RDIFF predicted noise change (PHW-configuration) to the measured production inlet test 

data was used to estimate the scarf inlet noise if the BPF and 2BPF tone increase did not occur. This is 

believed to be a valid flight condition prediction since it is believed that the flow distortion causing the 

tone increases will not be present with forward motion. It was concluded that without the BPF and 2BPF 

aggravation the 13° biplanar scarf shape is worth about 2 EPNdB and the triple layer lining is worth about 

1 EPNdB relative to a production inlet with lip lining and “stretched” to have the same total lining area as 

in the scarf inlet.

At cutback power it was concluded that the scarf shape significantly increases buzzsaw noise, but, 

except for the high frequencies (2 kHz and higher), the lining attenuated it to levels expected based on the 

measured treated production inlet levels. The RDIFF code however does not predict the attenuation of the 

buzz saw noise. The RDIFF calculated acoustic wave attenuation for the treated scarf inlet relative to the 

treated production inlet at the dominant buzz saw noise frequencies (<1 kHz) appeared to be reasonably 

close to the measured data. Besides buzz saw noise attenuation prediction, the other major miss by the 

RDIFF code was the prediction of the high frequency (2 kHz and higher) noise attenuation at the higher 

power conditions. Contrary to the RDIFF predictions, the measured high frequency attenuations were 

observed to fall off rapidly with increased power setting. It is speculated that this is due to a propagation 

effect through the high Mach number flow, which reduces the hardwall and treated inlet noise levels 

increasingly as engine power is increased while reducing the effect of the lining which is adjacent to the 

high Mach number flow. If this is true the RDIFF high frequency predictions may be valid for flight 

where the lip Mach numbers are much lower.

If the predicted high frequency attenuation is realized, the 13° biplanar scarf shape is worth about 

3 EPNdB and the triple layer lining is worth about 0.75 EPNdB relative to a production inlet with a 

treated lip and “stretched” to have lining area equal to the tested scarf inlet.

5.3 ICD Microphone Array

The ICD phased array can measure low-pressure compressor noise in the expected spinning modes.

The results for blade passage frequency are promising, but not completely satisfactory. This is probably 

due to the scattering of the modes by the nonuniform inlet and flow geometry.

At approach power, the inlet noise of the PW4098 has a strong contribution from the low pressure 

compressor. Blade passage tones and broadband fan noise are also important. High pressure compressor 

noise may play a role.

Blade passage frequency noise for the hardwall baseline inlet is dominated by the mode m = 17-18.

There is no obvious mode in the treated case.

The blade passage frequency noise of the treated scarf inlet is strongly concentrated around m = 5.

Beamforming images show a periodic source structure that may be consistent with generation of this 

mode.

Inlet lip lining is very effective in attenuating buzzsaw noise at low spinning orders.

The production lining is ineffective at takeoff power.

The production lining is largely ineffective for low-pressure compressor noise at approach power.
NASA/CR—2014-218088 115

5.4 BMT and Mode Array Tone Diagnostics

The mode array, in combination with the BMT’s and the flow visualization techniques, provided 

insight into some of the BPF noise-generation mechanisms for the scarf and production inlets. These data 

show the additional BPF noise for the scarf over the production inlet at subsonic speeds is due to a flow 

problem at the crown. The current scarf inlet, without a bellmouth, separates at the crown, which causes 

vortices to be shed into the fan. These vortices have long axial length scales, which are not stationary and 

hence, generate a non-deterministic BPF noise source in spinning orders that propagate out the nacelle. 

High peak Mach numbers around the inlet lip of the scarf inlet can be reduced with the use of a 

bellmouth, thus reducing the possibility of separated flow. Recent CFD calculations indicate that the 

separated flow in the scarf inlet will be eliminated with forward flight. Additional work is underway to 

determine the possibility of testing the scarf inlet with a bellmouth inlet.

The production inlet with the ICD did not show evidence of a non-deterministic BPF noise source in 

the mode array or BMT data. Since the production inlet uses a bellmouth inlet which reduces the Mach 

number near the lip, the probability of flow separation is minimized thereby eliminating the existence of 

these unsteady flow distortions. In addition, CFD results have shown the scarf inlet to have higher peak 

Mach numbers than the production inlet due its flow lines. Recent CFD runs have shown some evidence 

that the scarf inlet would not have the separation BPF noise source if it was in forward flight or had a 

bellmouth inlet. 

The cause of the BPF noise for the production inlet at subsonic speeds was never fully determined. 

The data was able to show that the production inlet BPF noise is due to a deterministic noise mechanism, 

but the source of the mechanism is not conclusive. The BMT data show steady distortions which exist in 

the boundary layer, yet it is not clear whether the noise generated from the interaction of the fan with 

these disturbances is sufficient to generate the observed BPF tones. The mode array data, however, shows 

that the noise source is spread out across all the propagating spinning orders. This implies that the source 

is not comprised of one interaction that can be determined and fixed. The source seems to be comprised 

of many propagating spinning orders. Mode measurements and/or BMT data during 

acceleration/deceleration and with the inlet rotated may provide the additional data to needed to help 

understand this noise source. These types of data were not taken during this test. 

The mode array data did not provide any evidence of large differences between the production and 

scarf deterministic noise sources. However, the BMT data clearly showed large steady distortions in the 

pressures measured on the blades for both inlets. There is not currently enough data to determine the 

amount of additional noise due to the steady distortion as measured with the BMT’s. The data clearly 

show that the unsteady part of the noise dominates the differences. Possibly, some of the fan-noise 

prediction codes could be used to predict the relative levels of the steady and unsteady noise sources from 

the BMT data to compare these sources. The data suggests that the steady source increase is insignificant 

in comparison. 

Clearly the mode array, BMT’s, and the visualization techniques are useful in helping understand 

complex BPF noise mechanisms. It is unlikely that the source of the BPF noise for the scarf inlet would 

have been discovered and so fully understood without these techniques. Likewise, there now exists a 

greater understanding of the type of source that needs to be eliminated in order to reduce subsonic BPF 

noise of current engines.

NASA/CR—2014-218088 117

Appendix

This appendix presents a complete listing of data records by Configuration and Run Number.

The convention (or key) to data record ID (database ID) is as follows:

RRRRRGRXXX

RRRRR = 5 digit run number. A unique run number is assigned for every configuration. A 

configuration represents a hardware or test set up change.

GR = Microphone type code where GR means a ground plane microphone as part of a radial array.

XXX = An index counter, a new index is given to each data point acquired within a run number.

SOURCE: Run 16078 CONFIG 1

PW4098 X841-10A,BOM HW INLET+FFCC,KEEL TO MICS,ICD,AFT WALLS,NO PT2 PRB

*** SORTED BY N1C ***

RECORD ID No. of ANGLES N1CORR N1OBS DATE TIME

16078GR017 24 0. 0. 111098 1257 

16078GR024 24 0. 0. 111198 0740 

16078GR003 24 571. 580. 110998 1321 

16078GR001 24 571. 580. 110998 1310 

16078GR002 24 571. 580. 110998 1314 

16078GR004 24 1499. 1521. 110998 1634 

16078GR033 24 1500. 1534. 111198 1221 

16078GR005 24 1503. 1526. 110998 1658 

16078GR006 24 1603. 1627. 110998 1702 

16078GR034 24 1605. 1641. 111198 1224 

16078GR035 24 1703. 1742. 111198 1227 

16078GR007 24 1704. 1729. 110998 1706 

16078GR008 24 1750. 1775. 110998 1710 

16078GR036 24 1754. 1794. 111198 1231 

16078GR009 24 1802. 1828. 110998 1713 

16078GR028 24 1804. 1844. 111198 1153 

16078GR010 24 1850. 1876. 110998 1721 

16078GR037 24 1854. 1895. 111198 1235 

16078GR011 24 1900. 1926. 110998 1725 

16078GR038 24 1902. 1944. 111198 1239 

16078GR012 24 2002. 2030. 110998 1728 

16078GR013 24 2102. 2130. 110998 1733 

16078GR014 24 2201. 2230. 110998 1738 

16078GR029 24 2204. 2252. 111198 1158 

16078GR046 24 2206. 2256. 111198 1311 

16078GR039 24 2250. 2301. 111198 1243 

16078GR015 24 2252. 2281. 110998 1743 

16078GR016 24 2300. 2329. 110998 1747 

16078GR040 24 2303. 2355. 111198 1246 

16078GR018 24 2351. 2392. 111098 1729 

16078GR041 24 2353. 2407. 111198 1249 

16078GR042 24 2403. 2457. 111198 1254 

16078GR019 24 2403. 2445. 111098 1735 

16078GR030 24 2451. 2506. 111198 1204 

16078GR020 24 2451. 2494. 111098 1738 

16078GR021 24 2452. 2493. 111098 1741 

16078GR043 24 2503. 2560. 111198 1257 

16078GR022 24 2503. 2545. 111098 1746 

16078GR023 24 2599. 2642. 111098 1751 

16078GR044 24 2602. 2662. 111198 1301 

16078GR027 24 2703. 2760. 111198 1119 

16078GR031 24 2711. 2771. 111198 1209 

16078GR045 24 2799. 2864. 111198 1305 

16078GR025 24 2803. 2864. 111198 1104 

16078GR032 24 2897. 2962. 111198 1214 

16078GR026 24 2906. 2967. 111198 1108 
NASA/CR—2014-218088 118

SOURCE: 16079 CONFIG 2

PW4098 X841-10A,BOM TRT INL +FFCC,KEEL TO MICS,ICD,AFT WALLS,NO PT2 PR

*** SORTED BY N1C ***

RECORD ID No. of ANGLES N1CORR N1OBS DATE TIME

16079GR001 24 0. 0. 111198 1732 

16079GR002 24 0. 0. 111198 1735 

16079GR003 24 566. 576. 111198 1748 

16079GR004 24 1479. 1504. 111198 1753 

16079GR042 24 1505. 1524. 111198 2044 

16079GR024 24 1506. 1527. 111198 1934 

16079GR025 24 1607. 1629. 111198 1937 

16079GR005 24 1607. 1634. 111198 1758 

16079GR006 24 1702. 1730. 111198 1801 

16079GR026 24 1705. 1728. 111198 1940 

16079GR007 24 1755. 1783. 111198 1804 

16079GR027 24 1759. 1781. 111198 1944 

16079GR028 24 1804. 1827. 111198 1946 

16079GR008 24 1806. 1835. 111198 1808 

16079GR029 24 1854. 1878. 111198 1950 

16079GR009 24 1862. 1891. 111198 1813 

16079GR030 24 1906. 1931. 111198 1952 

16079GR010 24 1907. 1938. 111198 1816 

16079GR011 24 2005. 2037. 111198 1819 

16079GR012 24 2106. 2137. 111198 1844 

16079GR031 24 2205. 2234. 111198 1955 

16079GR013 24 2207. 2240. 111198 1848 

16079GR032 24 2254. 2283. 111198 1958 

16079GR014 24 2262. 2295. 111198 1852 

16079GR015 24 2303. 2337. 111198 1854 

16079GR033 24 2307. 2338. 111198 2001 

16079GR034 24 2355. 2386. 111198 2004 

16079GR016 24 2358. 2392. 111198 1857 

16079GR035 24 2405. 2436. 111198 2007 

16079GR017 24 2408. 2443. 111198 1900 

16079GR036 24 2452. 2485. 111198 2010 

16079GR018 24 2454. 2489. 111198 1903 

16079GR037 24 2504. 2537. 111198 2013 

16079GR019 24 2506. 2542. 111198 1907 

16079GR038 24 2603. 2637. 111198 2016 

16079GR020 24 2605. 2642. 111198 1910 

16079GR039 24 2703. 2739. 111198 2019 

16079GR021 24 2704. 2743. 111198 1913 

16079GR022 24 2803. 2843. 111198 1916 

16079GR040 24 2804. 2840. 111198 2022 

16079GR041 24 2897. 2935. 111198 2026 

16079GR023 24 2904. 2946. 111198 1919 
NASA/CR—2014-218088 119

SOURCE: 16082 CONFIG 3

C3:SCARF HW INL,SIM HW BOM FFCC ,KEEL TO MICS,ICD,AFT WALLS,NO PT2 PR

*** SORTED BY N1C ***

RECORD ID No. of ANGLES N1CORR N1OBS DATE TIME

16082GR041 24 0. 0. 111798 0957 

16082GR042 24 0. 0. 111798 1000 

16082GR001 24 15. 15. 111698 0923 

16082GR002 24 568. 581. 111698 1020 

16082GR003 24 1503. 1534. 111698 1024 

16082GR023 24 1505. 1536. 111698 1138 

16082GR004 24 1596. 1629. 111698 1028 

16082GR024 24 1604. 1637. 111698 1142 

16082GR025 24 1705. 1740. 111698 1145 

16082GR005 24 1706. 1742. 111698 1031 

16082GR026 24 1753. 1789. 111698 1148 

16082GR006 24 1754. 1792. 111698 1035 

16082GR007 24 1804. 1843. 111698 1039 

16082GR027 24 1806. 1845. 111698 1151 

16082GR028 24 1854. 1893. 111698 1155 

16082GR008 24 1858. 1898. 111698 1042 

16082GR029 24 1900. 1940. 111698 1158 

16082GR009 24 1902. 1943. 111698 1046 

16082GR010 24 2003. 2046. 111698 1049 

16082GR011 24 2102. 2147. 111698 1053 

16082GR012 24 2202. 2249. 111698 1057 

16082GR030 24 2204. 2251. 111698 1202 

16082GR013 24 2252. 2299. 111698 1100 

16082GR031 24 2252. 2299. 111698 1205 

16082GR014 24 2302. 2351. 111698 1104 

16082GR032 24 2304. 2352. 111698 1208 

16082GR015 24 2351. 2401. 111698 1107 

16082GR033 24 2352. 2401. 111698 1211 

16082GR016 24 2403. 2455. 111698 1111 

16082GR034 24 2403. 2453. 111698 1214 

16082GR017 24 2454. 2506. 111698 1114 

16082GR035 24 2454. 2505. 111698 1217 

16082GR018 24 2500. 2553. 111698 1117 

16082GR036 24 2501. 2554. 111698 1222 

16082GR019 24 2600. 2656. 111698 1120 

16082GR037 24 2606. 2663. 111698 1225 

16082GR038 24 2701. 2758. 111698 1229 

16082GR020 24 2703. 2762. 111698 1124 

16082GR039 24 2803. 2862. 111698 1234 

16082GR021 24 2805. 2864. 111698 1128 

16082GR040 24 2901. 2964. 111698 1238 

16082GR022 24 2904. 2964. 111698 1132 
NASA/CR—2014-218088 120

SOURCE: 16083 CONFIG 12

C12:SCARF TRT INL W/HW LIP,HW FFCC,KEEL-MICS,ICD,AFT WALLS;NO PT PRB

*** SORTED BY N1C ***

RECORD ID No. of ANGLES N1CORR N1OBS DATE TIME

16083GR001 24 0. 0. 111798 1211 

16083GR002 24 0. 0. 111798 1213 

16083GR003 24 0. 0. 111798 1222 

16083GR004 24 552. 566. 111798 1237 

16083GR025 24 1499. 1534. 111798 1349 

16083GR005 24 1503. 1538. 111798 1240 

16083GR006 24 1598. 1636. 111798 1245 

16083GR026 24 1602. 1639. 111798 1353 

16083GR027 24 1704. 1743. 111798 1356 

16083GR007 24 1707. 1748. 111798 1248 

16083GR028 24 1755. 1796. 111798 1359 

16083GR008 24 1757. 1799. 111798 1251 

16083GR009 24 1804. 1847. 111798 1254 

16083GR029 24 1804. 1846. 111798 1402 

16083GR030 24 1852. 1895. 111798 1405 

16083GR010 24 1858. 1902. 111798 1257 

16083GR031 24 1902. 1947. 111798 1408 

16083GR011 24 1908. 1953. 111798 1301 

16083GR012 24 2006. 2053. 111798 1304 

16083GR013 24 2102. 2153. 111798 1307 

16083GR032 24 2200. 2253. 111798 1412 

16083GR014 24 2207. 2261. 111798 1310 

16083GR033 24 2253. 2307. 111798 1415 

16083GR015 24 2255. 2310. 111798 1314 

16083GR016 24 2301. 2360. 111798 1317 

16083GR034 24 2305. 2361. 111798 1418 

16083GR035 24 2351. 2410. 111798 1421 

16083GR017 24 2355. 2412. 111798 1321 

16083GR036 24 2401. 2460. 111798 1424 

16083GR018 24 2402. 2459. 111798 1324 

16083GR037 24 2450. 2510. 111798 1427 

16083GR019 24 2451. 2508. 111798 1327 

16083GR038 24 2501. 2564. 111798 1430 

16083GR020 24 2503. 2562. 111798 1331 

16083GR039 24 2602. 2667. 111798 1434 

16083GR021 24 2605. 2666. 111798 1334 

16083GR022 24 2701. 2764. 111798 1338 

16083GR040 24 2703. 2768. 111798 1437 

16083GR023 24 2803. 2869. 111798 1341 

16083GR041 24 2804. 2870. 111798 1440 

16083GR042 24 2903. 2973. 111798 1443 

16083GR024 24 2904. 2973. 111798 1344 
NASA/CR—2014-218088 121

SOURCE: 16084 CONFIG 14

C14:SCARF TRT INL W/TRT LIP,HW FFCC,KEEL-MICS,ICD,AFT WALLS;NO PT PRB

*** SORTED BY N1C ***

RECORD ID No. of ANGLES N1CORR N1OBS DATE TIME

16084GR001 24 0. 0. 111898 0733 

16084GR003 24 0. 0. 111898 0853 

16084GR043 24 0. 0. 111898 1247 

16084GR044 24 0. 0. 111898 1251 

16084GR002 24 1. 1. 111898 0735 

16084GR004 24 589. 601. 111898 0902 

16084GR005 24 1502. 1531. 111898 0907 

16084GR025 24 1514. 1547. 111898 1009 

16084GR006 24 1605. 1637. 111898 0910 

16084GR026 24 1607. 1643. 111898 1011 

16084GR027 24 1701. 1739. 111898 1014 

16084GR007 24 1705. 1739. 111898 0913 

16084GR008 24 1753. 1788. 111898 0916 

16084GR028 24 1757. 1795. 111898 1017 

16084GR009 24 1802. 1839. 111898 0920 

16084GR029 24 1805. 1843. 111898 1020 

16084GR034 24 1850. 1888. 111898 1038 

16084GR010 24 1851. 1889. 111898 0923 

16084GR011 24 1901. 1940. 111898 0926 

16084GR035 24 1904. 1943. 111898 1041 

16084GR012 24 2005. 2047. 111898 0929 

16084GR013 24 2103. 2146. 111898 0932 

16084GR030 24 2204. 2249. 111898 1025 

16084GR014 24 2208. 2254. 111898 0935 

16084GR015 24 2252. 2299. 111898 0938 

16084GR036 24 2254. 2300. 111898 1044 

16084GR037 24 2300. 2348. 111898 1047 

16084GR016 24 2302. 2353. 111898 0941 

16084GR017 24 2353. 2404. 111898 0944 

16084GR038 24 2353. 2403. 111898 1049 

16084GR039 24 2402. 2452. 111898 1052 

16084GR018 24 2404. 2456. 111898 0946 

16084GR031 24 2451. 2502. 111898 1028 

16084GR019 24 2452. 2503. 111898 0949 

16084GR020 24 2502. 2552. 111898 0953 

16084GR040 24 2502. 2553. 111898 1055 

16084GR021 24 2601. 2654. 111898 0956 

16084GR041 24 2602. 2658. 111898 1058 

16084GR032 24 2700. 2758. 111898 1031 

16084GR022 24 2704. 2761. 111898 0959 

16084GR023 24 2798. 2859. 111898 1002 

16084GR042 24 2799. 2862. 111898 1102 

16084GR033 24 2905. 2966. 111898 1034 

16084GR024 24 2907. 2969. 111898 1005 
NASA/CR—2014-218088 122

SOURCE: 16085 CONFIG 4

C04:SCARF TRT INL W/TRT LIP,TRT FFCC,KEEL-MICS,ICD,AFT WALLS;NO PT PRB

*** SORTED BY N1C ***

RECORD ID No. of ANGLES N1CORR N1OBS DATE TIME

16085GR001 24 0. 0. 111898 1316 

16085GR042 24 0. 0. 111898 1619 

16085GR043 24 0. 0. 111898 1624 

16085GR003 24 588. 602. 111898 1323 

16085GR024 24 1512. 1548. 111898 1431 

16085GR004 24 1515. 1549. 111898 1328 

16085GR005 24 1603. 1640. 111898 1331 

16085GR025 24 1608. 1646. 111898 1434 

16085GR026 24 1705. 1744. 111898 1437 

16085GR006 24 1706. 1746. 111898 1335 

16085GR027 24 1751. 1792. 111898 1440 

16085GR007 24 1757. 1799. 111898 1338 

16085GR008 24 1805. 1847. 111898 1341 

16085GR028 24 1805. 1848. 111898 1442 

16085GR009 24 1850. 1894. 111898 1344 

16085GR029 24 1854. 1899. 111898 1446 

16085GR002 24 1861. 1904. 111898 1320 

16085GR030 24 1899. 1945. 111898 1448 

16085GR010 24 1906. 1952. 111898 1347 

16085GR011 24 2005. 2053. 111898 1350 

16085GR012 24 2101. 2150. 111898 1353 

16085GR031 24 2202. 2257. 111898 1452 

16085GR013 24 2203. 2257. 111898 1356 

16085GR014 24 2252. 2306. 111898 1359 

16085GR032 24 2255. 2311. 111898 1455 

16085GR033 24 2304. 2357. 111898 1500 

16085GR015 24 2305. 2358. 111898 1403 

16085GR016 24 2352. 2404. 111898 1406 

16085GR034 24 2352. 2405. 111898 1503 

16085GR035 24 2401. 2456. 111898 1506 

16085GR017 24 2402. 2455. 111898 1408 

16085GR018 24 2451. 2505. 111898 1411 

16085GR036 24 2453. 2510. 111898 1509 

16085GR019 24 2504. 2561. 111898 1415 

16085GR037 24 2505. 2562. 111898 1512 

16085GR020 24 2601. 2662. 111898 1418 

16085GR038 24 2603. 2663. 111898 1515 

16085GR039 24 2701. 2763. 111898 1519 

16085GR021 24 2703. 2767. 111898 1421 

16085GR022 24 2802. 2870. 111898 1424 

16085GR040 24 2804. 2869. 111898 1522 

16085GR023 24 2901. 2972. 111898 1427 

16085GR041 24 2906. 2973. 111898 1528 
NASA/CR—2014-218088 123

SOURCE: 16086 CONFIG 5

C05:SCARF TRT INLET,R1 FFCC, KEEL-MICS,ICD,AFT WALLS;NO PT PROBE

*** SORTED BY N1C ***

RECORD ID No. of ANGLES N1CORR N1OBS DATE TIME

16086GR001 24 0. 0. 111998 1837 

16086GR002 24 0. 0. 111998 1841 

16086GR039 24 0. 0. 111998 2203 

16086GR040 24 0. 0. 111998 2207 

16086GR003 24 573. 586. 111998 1904 

16086GR004 24 1557. 1584. 111998 1909 

16086GR024 24 1557. 1584. 111998 2016 

16086GR005 24 1602. 1630. 111998 1913 

16086GR034 24 1611. 1636. 111998 2056 

16086GR035 24 1703. 1730. 111998 2058 

16086GR006 24 1704. 1734. 111998 1916 

16086GR007 24 1752. 1782. 111998 1920 

16086GR025 24 1753. 1782. 111998 2019 

16086GR026 24 1802. 1832. 111998 2023 

16086GR008 24 1803. 1834. 111998 1923 

16086GR009 24 1851. 1884. 111998 1927 

16086GR027 24 1854. 1884. 111998 2027 

16086GR010 24 1900. 1933. 111998 1930 

16086GR036 24 1904. 1934. 111998 2103 

16086GR011 24 2000. 2034. 111998 1933 

16086GR012 24 2106. 2142. 111998 1936 

16086GR028 24 2204. 2240. 111998 2031 

16086GR013 24 2204. 2242. 111998 1939 

16086GR014 24 2252. 2291. 111998 1943 

16086GR037 24 2255. 2290. 111998 2108 

16086GR015 24 2305. 2345. 111998 1946 

16086GR038 24 2307. 2343. 111998 2112 

16086GR016 24 2354. 2394. 111998 1949 

16086GR017 24 2401. 2443. 111998 1952 

16086GR029 24 2404. 2444. 111998 2035 

16086GR030 24 2450. 2490. 111998 2038 

16086GR018 24 2455. 2497. 111998 1955 

16086GR019 24 2501. 2544. 111998 1958 

16086GR031 24 2502. 2542. 111998 2041 

16086GR020 24 2601. 2646. 111998 2002 

16086GR032 24 2700. 2744. 111998 2045 

16086GR021 24 2702. 2748. 111998 2005 

16086GR022 24 2800. 2848. 111998 2008 

16086GR033 24 2900. 2946. 111998 2050 

16086GR023 24 2900. 2949. 111998 2011 
NASA/CR—2014-218088 124

SOURCE: 16087 CONFIG 6

CO6:SCARF TRT INLET,R2 FFCC, KEEL-MICS,ICD,AFT WALLS;NO PT PROBE

*** SORTED BY N1C ***

RECORD ID No. of ANGLES N1CORR N1OBS DATE TIME

16087GR001 24 0. 0. 112098 0800 

16087GR002 24 0. 0. 112098 0802 

16087GR042 24 0. 0. 112098 1355 

16087GR043 24 0. 0. 112098 1359 

16087GR003 24 565. 576. 112098 1044 

16087GR024 24 1554. 1588. 112098 1148 

16087GR004 24 1557. 1589. 112098 1047 

16087GR025 24 1602. 1639. 112098 1151 

16087GR005 24 1608. 1640. 112098 1050 

16087GR026 24 1700. 1739. 112098 1154 

16087GR006 24 1704. 1740. 112098 1053 

16087GR027 24 1750. 1789. 112098 1156 

16087GR007 24 1754. 1791. 112098 1056 

16087GR008 24 1802. 1841. 112098 1059 

16087GR028 24 1802. 1841. 112098 1159 

16087GR009 24 1851. 1892. 112098 1102 

16087GR029 24 1852. 1892. 112098 1203 

16087GR030 24 1902. 1942. 112098 1206 

16087GR010 24 1903. 1942. 112098 1105 

16087GR011 24 2002. 2044. 112098 1107 

16087GR012 24 2100. 2145. 112098 1110 

16087GR013 24 2200. 2248. 112098 1114 

16087GR031 24 2203. 2252. 112098 1209 

16087GR032 24 2251. 2303. 112098 1213 

16087GR014 24 2254. 2303. 112098 1117 

16087GR015 24 2302. 2353. 112098 1120 

16087GR033 24 2302. 2356. 112098 1215 

16087GR034 24 2353. 2408. 112098 1219 

16087GR016 24 2354. 2404. 112098 1123 

16087GR035 24 2400. 2456. 112098 1222 

16087GR017 24 2401. 2452. 112098 1125 

16087GR018 24 2450. 2502. 112098 1129 

16087GR036 24 2453. 2509. 112098 1224 

16087GR019 24 2501. 2557. 112098 1132 

16087GR037 24 2501. 2558. 112098 1228 

16087GR038 24 2603. 2661. 112098 1231 

16087GR020 24 2605. 2662. 112098 1135 

16087GR039 24 2699. 2762. 112098 1234 

16087GR021 24 2703. 2763. 112098 1138 

16087GR022 24 2802. 2863. 112098 1141 

16087GR040 24 2803. 2869. 112098 1237 

16087GR023 24 2905. 2972. 112098 1144 

16087GR041 24 2907. 2972. 112098 1239 
NASA/CR—2014-218088 125

SOURCE: 16088 CONFIG 7

C7:SCARF TRT INL,R2 8-3” HW SPL FFCC,KEEL TO MICS,ICD,AFT WALLS,NO PT PBE 

*** SORTED BY N1C ***

RECORD ID No. of ANGLES N1CORR N1OBS DATE TIME

16088GR001 24 0. 0. 112098 1652 

16088GR002 24 0. 0. 112098 1656 

16088GR003 24 0. 0. 112098 1711 

16088GR043 24 0. 0. 112098 2116 

16088GR044 24 0. 0. 112098 2129 

16088GR004 24 569. 579. 112098 1734 

16088GR005 24 1554. 1583. 112098 1740 

16088GR025 24 1557. 1584. 112098 1851 

16088GR006 24 1599. 1628. 112098 1743 

16088GR026 24 1602. 1629. 112098 1854 

16088GR007 24 1704. 1736. 112098 1746 

16088GR027 24 1707. 1735. 112098 1858 

16088GR028 24 1751. 1780. 112098 1903 

16088GR008 24 1759. 1791. 112098 1749 

16088GR009 24 1803. 1836. 112098 1752 

16088GR029 24 1803. 1834. 112098 1906 

16088GR030 24 1854. 1885. 112098 1912 

16088GR010 24 1856. 1890. 112098 1756 

16088GR031 24 1902. 1934. 112098 1915 

16088GR011 24 1903. 1938. 112098 1759 

16088GR012 24 2001. 2038. 112098 1801 

16088GR013 24 2102. 2140. 112098 1805 

16088GR032 24 2203. 2239. 112098 1919 

16088GR014 24 2206. 2246. 112098 1808 

16088GR015 24 2251. 2292. 112098 1811 

16088GR033 24 2253. 2290. 112098 1922 

16088GR016 24 2301. 2342. 112098 1814 

16088GR034 24 2303. 2341. 112098 1926 

16088GR017 24 2351. 2394. 112098 1816 

16088GR040 24 2354. 2392. 112098 1947 

16088GR018 24 2401. 2444. 112098 1819 

16088GR035 24 2402. 2441. 112098 1929 

16088GR036 24 2451. 2490. 112098 1932 

16088GR019 24 2452. 2495. 112098 1823 

16088GR037 24 2501. 2541. 112098 1935 

16088GR020 24 2503. 2547. 112098 1827 

16088GR021 24 2601. 2647. 112098 1831 

16088GR041 24 2607. 2648. 112098 1951 

16088GR022 24 2700. 2747. 112098 1836 

16088GR038 24 2700. 2744. 112098 1939 

16088GR023 24 2802. 2850. 112098 1839 

16088GR042 24 2803. 2847. 112098 1955 

16088GR024 24 2900. 2950. 112098 1845 

16088GR039 24 2903. 2949. 112098 1943 
NASA/CR—2014-218088 126

SOURCE: 16089 CONFIG 8

C8:SCARF TRT INL,R2 3-3” HW SPL FFCC,KEEL TO MICS,ICD,AFT WALLS,NO PT PBE 

*** SORTED BY N1C ***

RECORD ID No. of ANGLES N1CORR N1OBS DATE TIME

16089GR001 24 0. 0. 112198 0906 

16089GR044 24 0. 0. 112198 1159 

16089GR045 24 0. 0. 112198 1201 

16089GR046 24 0. 0. 112198 1205 

16089GR003 24 1. 1. 112198 0918 

16089GR002 24 1. 1. 112198 0909 

16089GR004 24 565. 575. 112198 0930 

16089GR005 24 1545. 1575. 112198 0933 

16089GR025 24 1551. 1584. 112198 1035 

16089GR006 24 1601. 1632. 112198 0936 

16089GR026 24 1607. 1641. 112198 1037 

16089GR027 24 1700. 1737. 112198 1040 

16089GR007 24 1702. 1735. 112198 0939 

16089GR028 24 1752. 1790. 112198 1043 

16089GR008 24 1752. 1786. 112198 0942 

16089GR029 24 1799. 1838. 112198 1046 

16089GR009 24 1804. 1840. 112198 0946 

16089GR030 24 1854. 1894. 112198 1050 

16089GR010 24 1857. 1893. 112198 0949 

16089GR031 24 1902. 1943. 112198 1053 

16089GR011 24 1903. 1941. 112198 0952 

16089GR012 24 2003. 2042. 112198 0955 

16089GR013 24 2100. 2142. 112198 0957 

16089GR014 24 2199. 2243. 112198 1001 

16089GR032 24 2201. 2249. 112198 1057 

16089GR033 24 2202. 2250. 112198 1101 

16089GR015 24 2252. 2298. 112198 1004 

16089GR034 24 2253. 2302. 112198 1105 

16089GR016 24 2303. 2350. 112198 1007 

16089GR035 24 2304. 2355. 112198 1108 

16089GR017 24 2352. 2400. 112198 1010 

16089GR036 24 2353. 2406. 112198 1110 

16089GR037 24 2401. 2455. 112198 1113 

16089GR018 24 2403. 2452. 112198 1013 

16089GR038 24 2453. 2508. 112198 1116 

16089GR019 24 2453. 2504. 112198 1016 

16089GR020 24 2501. 2553. 112198 1019 

16089GR039 24 2502. 2558. 112198 1120 

16089GR040 24 2604. 2661. 112198 1122 

16089GR021 24 2605. 2660. 112198 1022 

16089GR041 24 2700. 2761. 112198 1125 

16089GR022 24 2703. 2759. 112198 1024 

16089GR042 24 2802. 2865. 112198 1128 

16089GR023 24 2802. 2860. 112198 1027 

16089GR043 24 2903. 2969. 112198 1132 

16089GR024 24 2903. 2965. 112198 1030 
NASA/CR—2014-218088 127

SOURCE: 16090 CONFIG 9

C9:SCARF TRT INL,R2,3-1.25” HW SPL FFCC,KEEL-MICS,ICD,AFT WALLS,NO PT PBE 

*** SORTED BY N1C ***

RECORD ID No. of ANGLES N1CORR N1OBS DATE TIME

16090GR002 24 0. 0. 112198 1337 

16090GR003 24 0. 0. 112198 1341 

16090GR043 24 0. 0. 112198 2200 

16090GR044 24 0. 0. 112198 2205 

16090GR004 24 562. 576. 112198 1351 

16090GR005 24 1541. 1579. 112198 1355 

16090GR025 24 1546. 1582. 112198 1454 

16090GR026 24 1602. 1639. 112198 1457 

16090GR006 24 1605. 1644. 112198 1358 

16090GR007 24 1702. 1743. 112198 1400 

16090GR027 24 1702. 1742. 112198 1500 

16090GR008 24 1751. 1795. 112198 1403 

16090GR028 24 1754. 1796. 112198 1503 

16090GR009 24 1804. 1849. 112198 1406 

16090GR029 24 1807. 1850. 112198 1506 

16090GR010 24 1851. 1898. 112198 1409 

16090GR030 24 1854. 1900. 112198 1509 

16090GR011 24 1901. 1949. 112198 1412 

16090GR031 24 1901. 1949. 112198 1512 

16090GR012 24 2003. 2053. 112198 1415 

16090GR013 24 2104. 2156. 112198 1418 

16090GR014 24 2201. 2255. 112198 1420 

16090GR032 24 2205. 2261. 112198 1515 

16090GR033 24 2253. 2309. 112198 1518 

16090GR015 24 2257. 2311. 112198 1423 

16090GR016 24 2301. 2356. 112198 1426 

16090GR034 24 2302. 2360. 112198 1521 

16090GR017 24 2349. 2406. 112198 1429 

16090GR035 24 2353. 2411. 112198 1523 

16090GR018 24 2399. 2457. 112198 1431 

16090GR036 24 2403. 2462. 112198 1526 

16090GR019 24 2453. 2513. 112198 1434 

16090GR037 24 2457. 2516. 112198 1530 

16090GR020 24 2500. 2562. 112198 1437 

16090GR038 24 2502. 2562. 112198 1533 

16090GR039 24 2601. 2662. 112198 1535 

16090GR021 24 2605. 2669. 112198 1440 

16090GR040 24 2702. 2764. 112198 1539 

16090GR022 24 2703. 2770. 112198 1443 

16090GR041 24 2801. 2865. 112198 1542 

16090GR023 24 2804. 2872. 112198 1447 

16090GR024 24 2904. 2972. 112198 1450 

16090GR042 24 2907. 2973. 112198 1544 
NASA/CR—2014-218088 128

SOURCE: 16091 CONFIG 

C11:SCARF TRT INL,R2,3-0.625” HW SPL FFCC,KEEL-MICS,ICD,AFT WALLS,NO PT PB

*** SORTED BY N1C ***

RECORD ID No. of ANGLES N1CORR N1OBS DATE TIME

16091GR001 24 0. 0. 112198 1751 

16091GR002 24 0. 0. 112198 1755 

16091GR003 24 0. 0. 112198 1759 

16091GR043 24 0. 0. 112198 2132 

16091GR044 24 0. 0. 112198 2136 

16091GR004 24 555. 564. 112198 1808 

16091GR005 24 1547. 1574. 112198 1811 

16091GR030 24 1548. 1572. 112198 1931 

16091GR031 24 1607. 1631. 112198 1934 

16091GR006 24 1611. 1639. 112198 1814 

16091GR007 24 1703. 1733. 112198 1817 

16091GR032 24 1707. 1733. 112198 1936 

16091GR033 24 1749. 1777. 112198 1940 

16091GR008 24 1754. 1783. 112198 1820 

16091GR009 24 1801. 1832. 112198 1823 

16091GR014 24 1803. 1833. 112198 1843 

16091GR019 24 1852. 1883. 112198 1900 

16091GR034 24 1852. 1881. 112198 1943 

16091GR020 24 1901. 1932. 112198 1903 

16091GR035 24 1903. 1933. 112198 1946 

16091GR021 24 2007. 2040. 112198 1906 

16091GR022 24 2105. 2139. 112198 1908 

16091GR015 24 2201. 2237. 112198 1846 

16091GR010 24 2203. 2240. 112198 1826 

16091GR023 24 2252. 2288. 112198 1911 

16091GR036 24 2256. 2291. 112198 1950 

16091GR037 24 2303. 2338. 112198 1952 

16091GR024 24 2305. 2341. 112198 1914 

16091GR038 24 2351. 2386. 112198 1955 

16091GR025 24 2352. 2389. 112198 1916 

16091GR026 24 2401. 2438. 112198 1919 

16091GR039 24 2401. 2438. 112198 1957 

16091GR016 24 2450. 2490. 112198 1849 

16091GR011 24 2455. 2497. 112198 1830 

16091GR040 24 2500. 2538. 112198 2000 

16091GR027 24 2504. 2544. 112198 1921 

16091GR041 24 2601. 2640. 112198 2003 

16091GR028 24 2604. 2645. 112198 1924 

16091GR012 24 2702. 2749. 112198 1835 

16091GR017 24 2704. 2749. 112198 1853 

16091GR029 24 2801. 2845. 112198 1927 

16091GR042 24 2803. 2844. 112198 2006 

16091GR013 24 2902. 2951. 112198 1838 

16091GR018 24 2902. 2949. 112198 1856 
NASA/CR—2014-218088 129

SOURCE: 16092 CONFIG 10

C10:SCRF TRT INL,R2 3-1.25”+6”HW FFCC,KEEL-MICS,ICD,AFT WALL,NO PT PRB

*** SORTED BY N1C ***

RECORD ID No. of ANGLES N1CORR N1OBS DATE TIME

16092GR001 24 0. 0. 112298 0654 

16092GR002 24 0. 0. 112298 0733 

16092GR003 24 0. 0. 112298 0727 

16092GR044 24 0. 0. 112298 1358 

16092GR045 24 0. 0. 112298 1404 

16092GR043 24 568. 580. 112298 1350 

16092GR005 24 1549. 1582. 112298 1126 

16092GR025 24 1551. 1586. 112298 1231 

16092GR026 24 1602. 1638. 112298 1234 

16092GR006 24 1602. 1636. 112298 1129 

16092GR007 24 1702. 1738. 112298 1133 

16092GR027 24 1706. 1744. 112298 1237 

16092GR008 24 1750. 1788. 112298 1136 

16092GR028 24 1752. 1791. 112298 1244 

16092GR009 24 1803. 1842. 112298 1139 

16092GR029 24 1805. 1846. 112298 1247 

16092GR030 24 1852. 1894. 112298 1253 

16092GR010 24 1852. 1891. 112298 1142 

16092GR031 24 1902. 1946. 112298 1256 

16092GR011 24 1902. 1944. 112298 1145 

16092GR012 24 2002. 2048. 112298 1148 

16092GR013 24 2104. 2152. 112298 1152 

16092GR014 24 2202. 2251. 112298 1156 

16092GR032 24 2203. 2253. 112298 1301 

16092GR015 24 2250. 2298. 112298 1159 

16092GR033 24 2251. 2301. 112298 1304 

16092GR016 24 2301. 2351. 112298 1203 

16092GR034 24 2302. 2354. 112298 1306 

16092GR035 24 2353. 2406. 112298 1309 

16092GR017 24 2356. 2405. 112298 1206 

16092GR018 24 2400. 2454. 112298 1209 

16092GR036 24 2401. 2456. 112298 1312 

16092GR019 24 2450. 2506. 112298 1212 

16092GR037 24 2452. 2509. 112298 1315 

16092GR038 24 2501. 2558. 112298 1319 

16092GR020 24 2501. 2556. 112298 1215 

16092GR021 24 2603. 2660. 112298 1218 

16092GR039 24 2604. 2663. 112298 1323 

16092GR040 24 2701. 2763. 112298 1326 

16092GR022 24 2702. 2761. 112298 1221 

16092GR023 24 2800. 2864. 112298 1224 

16092GR041 24 2802. 2864. 112298 1332 

16092GR024 24 2902. 2968. 112298 1227 

16092GR042 24 2905. 2970. 112298 1334 
NASA/CR—2014-218088 130

SOURCE: 16093 CONFIG 6A

C06:SCARF TRT INL,w/TRIPS,FFCC R2,KEEL-MICS,ICD,AFT WALLS;NO PT

*** SORTED BY N1C ***

RECORD ID No. of ANGLES N1CORR N1OBS DATE TIME

16093GR001 24 0. 0. 112398 0906 

16093GR002 24 0. 0. 112398 0909 

16093GR024 24 0. 0. 112398 1405 

16093GR025 24 0. 0. 112398 1411 

16093GR003 24 558. 570. 112398 1227 

16093GR004 24 1551. 1586. 112398 1230 

16093GR005 24 1606. 1641. 112398 1233 

16093GR006 24 1702. 1738. 112398 1236 

16093GR007 24 1756. 1794. 112398 1239 

16093GR008 24 1802. 1840. 112398 1242 

16093GR009 24 1851. 1888. 112398 1245 

16093GR010 24 1903. 1940. 112398 1248 

16093GR011 24 2000. 2040. 112398 1251 

16093GR012 24 2103. 2146. 112398 1254 

16093GR013 24 2201. 2246. 112398 1258 

16093GR014 24 2252. 2298. 112398 1300 

16093GR015 24 2302. 2349. 112398 1303 

16093GR016 24 2350. 2399. 112398 1306 

16093GR017 24 2398. 2450. 112398 1309 

16093GR018 24 2450. 2503. 112398 1312 

16093GR019 24 2501. 2556. 112398 1316 

16093GR020 24 2604. 2659. 112398 1319 

16093GR021 24 2702. 2762. 112398 1321 

16093GR022 24 2801. 2860. 112398 1324 

16093GR023 24 2904. 2964. 112398 1327 
NASA/CR—2014-218088 131

SOURCE: 16094 CONFIG 10A

C10A:SCRF INL W/TRIPS,R2 3-1.25”+6”HW FFCC,KEEL-MICS,ICD,AFTWALL,NO PT

*** SORTED BY N1C ***

RECORD ID No. of ANGLES N1CORR N1OBS DATE TIME

16094GR001 24 0. 0. 112398 1610 

16094GR002 24 0. 0. 112398 1613 

16094GR003 24 0. 0. 112498 1227 

16094GR044 24 0. 0. 112498 1210 

16094GR045 24 0. 0. 112498 1213 

16094GR046 24 0. 0. 112498 1216 

16094GR004 24 587. 596. 112498 1256 

16094GR005 24 1546. 1568. 112498 1300 

16094GR025 24 1550. 1574. 112498 1103 

16094GR026 24 1602. 1627. 112498 1106 

16094GR006 24 1607. 1631. 112498 1005 

16094GR027 24 1701. 1728. 112498 1109 

16094GR007 24 1702. 1727. 112498 1008 

16094GR028 24 1749. 1777. 112498 1112 

16094GR008 24 1753. 1780. 112498 1011 

16094GR009 24 1801. 1828. 112498 1014 

16094GR029 24 1802. 1831. 112498 1114 

16094GR030 24 1849. 1881. 112498 1117 

16094GR010 24 1851. 1880. 112498 1018 

16094GR031 24 1903. 1935. 112498 1120 

16094GR011 24 1903. 1931. 112498 1021 

16094GR012 24 2004. 2034. 112498 1024 

16094GR043 24 2100. 2139. 112498 1207 

16094GR013 24 2104. 2135. 112498 1027 

16094GR014 24 2203. 2237. 112498 1029 

16094GR032 24 2203. 2240. 112498 1123 

16094GR033 24 2251. 2291. 112498 1126 

16094GR015 24 2254. 2288. 112498 1033 

16094GR016 24 2301. 2336. 112498 1035 

16094GR034 24 2301. 2342. 112498 1129 

16094GR017 24 2352. 2387. 112498 1038 

16094GR035 24 2353. 2394. 112498 1132 

16094GR036 24 2400. 2442. 112498 1134 

16094GR018 24 2401. 2437. 112498 1041 

16094GR019 24 2452. 2489. 112498 1044 

16094GR037 24 2453. 2496. 112498 1137 

16094GR020 24 2503. 2542. 112498 1047 

16094GR038 24 2503. 2546. 112498 1140 

16094GR039 24 2599. 2644. 112498 1143 

16094GR021 24 2600. 2640. 112498 1049 

16094GR022 24 2701. 2742. 112498 1052 

16094GR040 24 2701. 2749. 112498 1146 

16094GR023 24 2802. 2845. 112498 1056 

16094GR041 24 2802. 2852. 112498 1149 

16094GR042 24 2901. 2955. 112498 1152 

16094GR024 24 2906. 2952. 112498 1059 

END OF PHASE 1A TEST CONFIGURATIONS AND ENGINE X841-10A
NASA/CR—2014-218088 132

START OF PHASE 1B TEST CONFIGURATIONS AND ENGINE X841-12

SOURCE: 16101 CONFIG P3

P3-BOM ROT INLET/BM,TRT F&A FC,BMTs,KULITES,ICD,NO PT2 PROBE

*** SORTED BY N1C ***

RECORD ID No. of ANGLES N1CORR N1OBS DATE TIME

16101GR001 22 0. 0. 060999 1445 

16101GR002 22 0. 0. 060999 1447 

16101GR003 22 589. 605. 060999 1522 

16101GR004 22 1616. 1659. 060999 1526 

16101GR016 22 1702. 1736. 060999 2117 

16101GR005 22 1709. 1753. 060999 1531 

16101GR015 22 1805. 1841. 060999 2114 

16101GR006 22 1809. 1855. 060999 1535 

16101GR007 22 1902. 1950. 060999 1538 

16101GR008 22 2008. 2059. 060999 1542 

16101GR009 22 2208. 2263. 060999 1546 

16101GR010 22 2303. 2348. 060999 2058 

16101GR011 22 2405. 2452. 060999 2101 

16101GR012 22 2451. 2498. 060999 2104 

16101GR013 22 2603. 2653. 060999 2107 

16101GR014 22 2805. 2859. 060999 2110 

SOURCE: 16102 CONFIG P3

P3-BOM ROT INLET/BM,TRT F&A FC,BMTs,KULITES,NO FLOOR ICD,NO PT2 PROBE

*** SORTED BY N1C ***

RECORD ID No. of ANGLES N1CORR N1OBS DATE TIME

16102GR001 22 0. 0. 061099 0921 

16102GR002 22 0. 0. 061099 0932 

16102GR003 22 0. 0. 061199 0919 

16102GR004 22 587. 604. 061199 1232 

16102GR019 22 1619. 1660. 061199 1344 

16102GR005 22 1627. 1672. 061199 1236 

16102GR006 22 1701. 1747. 061199 1239 

16102GR007 22 1800. 1851. 061199 1242 

16102GR008 22 1904. 1958. 061199 1245 

16102GR009 22 2003. 2060. 061199 1248 

16102GR010 22 2202. 2262. 061199 1251 

16102GR011 22 2300. 2364. 061199 1254 

16102GR012 22 2401. 2470. 061199 1257 

16102GR013 22 2451. 2521. 061199 1300 

16102GR015 22 2502. 2573. 061199 1319 

16102GR014 22 2603. 2679. 061199 1305 

16102GR016 22 2703. 2778. 061199 1323 

16102GR018 22 2802. 2874. 061199 1334 

16102GR017 22 2805. 2879. 061199 1325 
NASA/CR—2014-218088 133

SOURCE: 16103 CONFIG S1

S1-SCARF ROT INLET,TRT-F&A FCC,BMTs,KULITES,NO FLOOR ICD,NO PT2 PROBE

*** SORTED BY N1C ***

RECORD ID No. of ANGLES N1CORR N1OBS DATE TIME

16103GR001 22 0. 0. 061599 1006 

16103GR002 22 0. 0. 061599 1013 

16103GR003 22 0. 0. 061599 1042 

16103GR004 22 0. 0. 061599 1048 

16103GR005 22 586. 602. 061599 1323 

16103GR021 22 1624. 1667. 061599 1432 

16103GR006 22 1625. 1666. 061599 1328 

16103GR007 22 1704. 1744. 061599 1331 

16103GR008 22 1804. 1846. 061599 1334 

16103GR022 22 1804. 1852. 061599 1438 

16103GR009 22 1903. 1947. 061599 1337 

16103GR010 22 2001. 2047. 061599 1341 

16103GR023 22 2001. 2055. 061599 1442 

16103GR011 22 2203. 2253. 061599 1344 

16103GR024 22 2203. 2262. 061599 1445 

16103GR012 22 2306. 2359. 061599 1347 

16103GR013 22 2400. 2455. 061599 1350 

16103GR025 22 2451. 2518. 061599 1449 

16103GR014 22 2454. 2511. 061599 1353 

16103GR015 22 2504. 2562. 061599 1356 

16103GR016 22 2600. 2662. 061599 1400 

16103GR019 22 2601. 2668. 061599 1423 

16103GR017 22 2702. 2772. 061599 1414 

16103GR020 22 2799. 2874. 061599 1426 

16103GR018 22 2803. 2876. 061599 1418 

SOURCE: 16104 CONFIG S3

S3-Scarf ROT inlet,Trt F&A FC,BMTs,kulites,ICD no flr,no pt2,flo viz

*** SORTED BY N1C ***

RECORD ID No. of ANGLES N1CORR N1OBS DATE TIME

16104GR001 22 0. 0. 061599 2202 

16104GR006 22 209. 213. 061699 1007 

16104GR002 22 586. 597. 061599 2234 

16104GR011 22 1616. 1656. 061699 1211 

16104GR007 22 1623. 1662. 061699 1017 

16104GR003 22 1623. 1653. 061599 2241 

16104GR008 22 2001. 2051. 061699 1022 

16104GR012 22 2001. 2054. 061699 1215 

16104GR004 22 2002. 2038. 061599 2246 

16104GR014 22 2401. 2462. 061699 1221 

16104GR005 22 2404. 2447. 061599 2250 

16104GR009 22 2407. 2466. 061699 1025 

16104GR010 22 2800. 2869. 061699 1029 

16104GR013 22 2808. 2880. 061699 1218 
NASA/CR—2014-218088 134

SOURCE: 16105 CONFIG S13A

S13A-Scarf ROT HW BOT 90deg diff+lip,TRT-FCC,BMTs,kulites,ICD no gp,no Pt2

*** SORTED BY N1C ***

RECORD ID No. of ANGLES N1CORR N1OBS DATE TIME

16105GR001 22 0. 0. 061799 1148 

16105GR002 22 0. 0. 061799 1154 

16105GR003 22 0. 0. 061799 1203 

16105GR004 22 587. 602. 061799 1229 

16105GR018 22 1622. 1662. 061799 1350 

16105GR005 22 1623. 1665. 061799 1232 

16105GR006 22 1704. 1748. 061799 1235 

16105GR007 22 1802. 1848. 061799 1238 

16105GR019 22 1806. 1852. 061799 1353 

16105GR008 22 1904. 1954. 061799 1241 

16105GR020 22 2000. 2050. 061799 1356 

16105GR009 22 2001. 2047. 061799 1323 

16105GR021 22 2200. 2255. 061799 1400 

16105GR010 22 2203. 2252. 061799 1326 

16105GR011 22 2303. 2353. 061799 1329 

16105GR012 22 2406. 2460. 061799 1332 

16105GR013 22 2452. 2508. 061799 1334 

16105GR022 22 2453. 2517. 061799 1404 

16105GR014 22 2502. 2560. 061799 1337 

16105GR015 22 2603. 2664. 061799 1340 

16105GR016 22 2699. 2766. 061799 1343 

16105GR017 22 2807. 2875. 061799 1345 

SOURCE: 16106 CONFIG S15A-1

S15A-1 Scarf ROT 3” circ HW fwd of “A” flg, Trt FCC, BMT&KUL, ICD, no pb

*** SORTED BY N1C ***

RECORD ID No. of ANGLES N1CORR N1OBS DATE TIME

16106GR001 22 0. 0. 061899 0845 

16106GR002 22 0. 0. 061899 0852 

16106GR003 22 0. 0. 062199 0741 

16106GR004 22 0. 0. 062199 0745 

16106GR005 22 0. 0. 062199 0924 

16106GR006 22 585. 597. 062199 0933 

16106GR007 22 1626. 1662. 062199 0939 

16106GR020 22 1626. 1662. 062199 1018 

16106GR008 22 1706. 1744. 062199 0942 

16106GR009 22 1801. 1841. 062199 0945 

16106GR021 22 1801. 1842. 062199 1021 

16106GR010 22 1901. 1943. 062199 0948 

16106GR022 22 1998. 2044. 062199 1024 

16106GR011 22 2003. 2049. 062199 0950 

16106GR023 22 2169. 2219. 062199 1027 

16106GR012 22 2201. 2252. 062199 0953 

16106GR027 22 2201. 2253. 062199 1041 

16106GR013 22 2302. 2355. 062199 0956 

16106GR014 22 2404. 2459. 062199 0959 

16106GR024 22 2449. 2506. 062199 1030 

16106GR015 22 2453. 2509. 062199 1001 

16106GR016 22 2500. 2558. 062199 1004 

16106GR025 22 2600. 2661. 062199 1033 

16106GR017 22 2602. 2660. 062199 1007 

16106GR018 22 2704. 2763. 062199 1010 

16106GR019 22 2802. 2863. 062199 1012 

16106GR026 22 2805. 2871. 062199 1037 
NASA/CR—2014-218088 135

SOURCE: 16107 CONFIG S15A-3

S15A-3 Scarf ROT 3-3”x 6” HW strips fwd of “A” Trt FCC,BMT&KUL, ICD, no pb

*** SORTED BY N1C ***

RECORD ID No. of ANGLES N1CORR N1OBS DATE TIME

16107GR001 22 0. 0. 062199 1347 

16107GR002 22 0. 0. 062199 1354 

16107GR003 22 0. 0. 062199 1357 

16107GR004 22 584. 600. 062199 1437 

16107GR005 22 1623. 1670. 062199 1443 

16107GR006 22 1703. 1750. 062199 1446 

16107GR007 22 1804. 1852. 062199 1448 

16107GR008 22 1905. 1957. 062199 1451 

16107GR009 22 2005. 2060. 062199 1454 

16107GR010 22 2199. 2263. 062199 1457 

16107GR011 22 2300. 2368. 062199 1500 

16107GR012 22 2401. 2470. 062199 1503 

16107GR013 22 2452. 2522. 062199 1505 

16107GR017 22 2452. 2521. 062199 1518 

16107GR014 22 2501. 2574. 062199 1508 

16107GR015 22 2605. 2677. 062199 1512 

16107GR016 22 2703. 2778. 062199 1514 

SOURCE: 16108 CONFIG S15A-4

S15A-4 Scarf ROT,HW Bott 90 deg Throat&Lip,TRT FCC, BMT&KUL, ICD, no pb

*** SORTED BY N1C ***

RECORD ID No. of ANGLES N1CORR N1OBS DATE TIME

16108GR001 22 0. 0. 062199 2246 

16108GR002 22 0. 0. 062199 2249 

16108GR024 22 0. 0. 062299 0139 

16108GR025 22 0. 0. 062299 0142 

16108GR003 22 587. 596. 062199 2354 

16108GR005 22 1615. 1639. 062299 0007 

16108GR017 22 1627. 1653. 062299 0045 

16108GR004 22 1696. 1722. 062299 0002 

16108GR006 22 1803. 1831. 062299 0010 

16108GR018 22 1806. 1835. 062299 0048 

16108GR007 22 1901. 1930. 062299 0013 

16108GR008 22 1999. 2030. 062299 0016 

16108GR019 22 2000. 2031. 062299 0051 

16108GR020 22 2202. 2237. 062299 0054 

16108GR009 22 2204. 2239. 062299 0020 

16108GR010 22 2304. 2340. 062299 0023 

16108GR011 22 2399. 2436. 062299 0026 

16108GR012 22 2451. 2489. 062299 0028 

16108GR021 22 2451. 2489. 062299 0059 

16108GR013 22 2503. 2542. 062299 0031 

16108GR022 22 2600. 2640. 062299 0102 

16108GR014 22 2605. 2646. 062299 0034 

16108GR015 22 2713. 2755. 062299 0037 

16108GR016 22 2801. 2845. 062299 0041 

16108GR023 22 2806. 2849. 062299 0105 
NASA/CR—2014-218088 136

SOURCE: 16109 CONFIG S16

S16-Scarf TRT,keel dwn ,TRT FCC, BMTs,Kulites,ICD no gp, no PT2 probe

*** SORTED BY N1C ***

RECORD ID No. of ANGLES N1CORR N1OBS DATE TIME

16109GR001 24 0. 0. 062299 0920 

16109GR002 24 0. 0. 062299 0926 

16109GR003 24 0. 0. 062299 0932 

16109GR004 24 0. 0. 062399 0712 

16109GR005 24 0. 0. 062399 0902 

16109GR006 24 585. 599. 062399 0958 

16109GR020 24 1623. 1664. 062399 1046 

16109GR007 24 1625. 1664. 062399 1004 

16109GR008 24 1702. 1743. 062399 1007 

16109GR009 24 1799. 1841. 062399 1010 

16109GR021 24 1803. 1849. 062399 1049 

16109GR027 24 1804. 1858. 062399 1351 

16109GR010 24 1909. 1956. 062399 1013 

16109GR011 24 2000. 2050. 062399 1016 

16109GR022 24 2002. 2053. 062399 1053 

16109GR023 24 2201. 2257. 062399 1056 

16109GR012 24 2203. 2256. 062399 1019 

16109GR028 24 2208. 2275. 062399 1354 

16109GR013 24 2299. 2354. 062399 1022 

16109GR014 24 2405. 2463. 062399 1025 

16109GR015 24 2449. 2510. 062399 1028 

16109GR024 24 2454. 2518. 062399 1059 

16109GR029 24 2457. 2525. 062399 1357 

16109GR016 24 2500. 2563. 062399 1030 

16109GR017 24 2601. 2668. 062399 1033 

16109GR025 24 2602. 2668. 062399 1102 

16109GR030 24 2603. 2673. 062399 1400 

16109GR018 24 2701. 2771. 062399 1036 

16109GR019 24 2799. 2871. 062399 1041 

16109GR026 24 2803. 2873. 062399 1105 

SOURCE: 16119 CONFIG S16A

S16-Scarf TRT,keel dwn ,TRT FCC, BMTs,Kulites,ICD -12”,no GP,no PT2 Pb

*** SORTED BY N1C ***

RECORD ID No. of ANGLES N1CORR N1OBS DATE TIME

16110GR001 24 0. 0. 062399 1208 

16110GR002 24 586. 602. 062399 1217 

16110GR003 24 1623. 1667. 062399 1222 

16110GR004 24 1803. 1853. 062399 1225 

16110GR005 24 2001. 2056. 062399 1228 

16110GR006 24 2202. 2262. 062399 1232 

16110GR007 24 2455. 2523. 062399 1235 

16110GR008 24 2605. 2678. 062399 1238 

16110GR009 24 2801. 2880. 062399 1241 
NASA/CR—2014-218088 137

SOURCE: 16111 CONFIG S16B

S16-Scarf TRT,keel dwn ,TRT FCC, BMTs,Kulites,ICD +12”,no GP,no PT2 Pb

*** SORTED BY N1C ***

RECORD ID No. of ANGLES N1CORR N1OBS DATE TIME

16111GR001 24 0. 0. 062399 1546 

16111GR002 24 0. 0. 062399 1551 

16111GR011 24 0. 0. 062399 1744 

16111GR012 24 0. 0. 062399 1748 

16111GR003 24 587. 601. 062399 1644 

16111GR004 24 1616. 1655. 062399 1648 

16111GR005 24 1802. 1846. 062399 1651 

16111GR006 24 2001. 2050. 062399 1655 

16111GR007 24 2201. 2256. 062399 1658 

16111GR008 24 2452. 2512. 062399 1701 

16111GR009 24 2601. 2664. 062399 1704 

16111GR010 24 2802. 2869. 062399 1707 

SOURCE: 16112 CONFIG S13A-1

S13A-1 Scarf Rot HW bot 90 deg diff, TRT LIP&FCC, BMTs,ICD no GP,no PT2 Pb

*** SORTED BY N1C ***

RECORD ID No. of ANGLES N1CORR N1OBS DATE TIME

16112GR001 24 0. 0. 062499 1850 

16112GR002 24 0. 0. 062499 1853 

16112GR024 24 0. 0. 062499 2327 

16112GR003 24 588. 598. 062499 2148 

16112GR017 24 1624. 1652. 062499 2234 

16112GR004 24 1624. 1651. 062499 2151 

16112GR005 24 1703. 1732. 062499 2154 

16112GR018 24 1803. 1835. 062499 2237 

16112GR006 24 1803. 1834. 062499 2157 

16112GR007 24 1904. 1937. 062499 2201 

16112GR019 24 2002. 2038. 062499 2240 

16112GR008 24 2004. 2038. 062499 2204 

16112GR020 24 2201. 2241. 062499 2244 

16112GR009 24 2201. 2239. 062499 2207 

16112GR010 24 2301. 2341. 062499 2210 

16112GR011 24 2403. 2445. 062499 2213 

16112GR012 24 2453. 2496. 062499 2216 

16112GR021 24 2455. 2498. 062499 2247 

16112GR013 24 2505. 2549. 062499 2219 

16112GR022 24 2604. 2649. 062499 2251 

16112GR014 24 2605. 2651. 062499 2222 

16112GR015 24 2702. 2750. 062499 2225 

16112GR016 24 2802. 2851. 062499 2228 

16112GR023 24 2804. 2854. 062499 2254 
NASA/CR—2014-218088 138

SOURCE: 16113 CONFIG S17

S17-Scarf Rot, TRT Inlet+FCC, BMT’s, ICD no GP, no walls, PT2 Probe in

*** SORTED BY N1C ***

RECORD ID No. of ANGLES N1CORR N1OBS DATE TIME

16113GR001 32 0. 0. 062599 0947 

16113GR002 32 0. 0. 062599 1112 

16113GR003 32 0. 0. 062599 1115 

16113GR004 32 584. 599. 062599 1247 

16113GR005 32 1618. 1661. 062599 1252 

16113GR018 32 1620. 1658. 062599 1348 

16113GR006 32 1702. 1747. 062599 1256 

16113GR019 32 1800. 1842. 062599 1352 

16113GR007 32 1801. 1850. 062599 1300 

16113GR008 32 1903. 1955. 062599 1304 

16113GR009 32 2002. 2057. 062599 1308 

16113GR020 32 2004. 2047. 062599 1356 

16113GR010 32 2204. 2264. 062599 1312 

16113GR011 32 2307. 2369. 062599 1316 

16113GR012 32 2401. 2466. 062599 1320 

16113GR013 32 2451. 2520. 062599 1323 

16113GR021 32 2453. 2503. 062599 1400 

16113GR014 32 2502. 2570. 062599 1328 

16113GR015 32 2601. 2671. 062599 1332 

16113GR016 32 2709. 2773. 062599 1336 

16113GR017 32 2801. 2867. 062599 1343 

16113GR022 32 2809. 2864. 062599 1404 
NASA/CR—2014-218088 139

References

1. Dougherty, R. P., “Nacelle Acoustic Design by Ray Tracing in Three Dimensions”, AIAA 96-1773, 

2nd AIAA /CEAS Aeroacoustics Conference, May 6-8, 1996/State College, PA.

2. Dougherty, R. P. and Stoker, R. W. “Sidelobe Suppression for Phased Array Aeroacoustic 

Measurements,” AIAA/CEAS 98-2242

3. J. M. Tyler and T. G. Sofrin, “Axial Flow Compressor Noise,” SAE Trans. v. 70, pp. 309-332, 1962

4. Ganz, Joppa, Patten, and Scharpf, “Boeing 18-Inch Fan Rig Broadband Noise Test,” Boeing 

Commercial Airplane Group, Seattle, WA, NASA/CR—1998-208704.
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4. TITLE AND SUBTITLE

Pratt & Whitney/Boeing Engine Validation of Noise Reduction Concepts 

Final Report for NASA Contract NAS3-97144, Phase 1

5a. CONTRACT NUMBER

NAS3-97144 

5b. GRANT NUMBER

5c. PROGRAM ELEMENT NUMBER

6. AUTHOR(S)

Mathews, Douglas, C.; Bock, Larry, A.; Bielak, Gerald, W.; Dougherty, R., P.; Premo, John, 

W.; Scharpf, Dan, F.; Yu, Jia 

5d. PROJECT NUMBER

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WBS 473452.02.03.07.06.01.02 

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United Technologies Corporation, Pratt & Whitney 

400 Main Street 

East Hartford, Connecticut 06108 

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This report presents results of the work completed in Phase 1 of the Engine Validation of Noise Reduction Concepts (EVNRC) contract. The 

purpose of the program was to validate, through engine testing, advanced noise reduction concepts aimed at reducing engine noise up to 6 

EPNdB and improving nacelle suppression by 50 percent relative to 1992 technology. Phase 2 of the program is currently near completion 

and upon its conclusion will be summarized in a separate report. 

15. SUBJECT TERMS

Fans; Noise; Inlet aerodynamics 

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