UHB Demonstrator Interior Noise Control Flight Tests and ...€¦ · intensity survey, and Robert...

NASA Contractor Report 181897

UHB Demonstrator

Interior Noise Control

Flight Tests and Analysis

M.A. Simpson, P.M. Druez, A.J. Kimbrough,

M.P. Brock, P.L. Burg6, G.P. Mathur,

M.R. Cannon, and B.N. Tran

DOUGLAS AIRCRAFT COMPANY

MCDONNELL DOUGLAS CORPORATION

LONG BEACH, CA 90846

CONTRACT NAS1-18037

OCTOBER 1989

N/ ANat_ona, Aeronautics and

qSpace ACmLn_TatLon

Langley Research Center

Hampton,Virginia 23665-5225 ....

(NASA-CR-18189T) UHi5 DEMONSIRATOR INTERIOR

NOISE C._NTRF)L FLIGHT TtSTS AND ANALYSIS

Final Report (Douglas Aircraft Co.) 1_8 pCSCL 20A

G3/71

N.o0-13198

Uncl as

0239_79

https://ntrs.nasa.gov/search.jsp?R=19900003882 2020-07-26T10:14:09+00:00Z

NASA Contractor Report 181897

UHB Demonstrator

Interior Noise Control

Flight Tests and Analysis

M.A. Simpson, P.M. Druez, A.J. Kimbrough,

M.P. Brock, P.L. Burg6, G.P. Mathur,

M.R. Cannon, and B.N. Tran

DOUGLAS AIRCRAFT COMPANY

MCDONNELL DOUGLAS CORPORATION

LONG BEACH, CA 90846

CONTRACT NAS1-18037

OCTOBER 1989

Nalional Aoronaut_cs and

,%,ha(:(_Adrnmislralion

Langley Research CenterHampton, Virginia 23665-5225

Preface

This report was prepared by McDonnell Douglas Corporation under Task Assign-

meut 4 ()f Contract NAS1-18()37 with NASA l, aagley Research (',enter. The NASA

t(-chnical monitor was Dr. Kevin P. Shepherd.

Several individuals at Douglas Aircraft Company made significant contributions

to this study. Craig van de Lune was responsible for acoustic and vibration data

acquisition on the Demonstrator. Dr. Mark Lang planned and conducted the sound

intensity survey, and Robert Boyd analyzed survey data. Dr. Daryl May designed the

Quiet Cabin noise control treatment package, and Dr. Mahendra Joshi reviewed the

report and offered many constructive comments throughout this program.

PRECEDING PAGE BLAhlK I'lL, i" FILMED

iii

Contents

3

4

5

Introduction 1

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

1.2 Objectives and Approach .......................... 3

1.3 Report Overview ............................... 4

Flight Test Program 6

2.1 Test Aircraft ................................. 6

2.2 Propulsion System .............................. 6

2.3 Interior Noise Test Points .......................... 7

Data Acquisition and Analysis 18

3.1 Acquisition System .............................. 18

3.2 Transducers .................................. 18

3.3 Data Processing and Analysis ........................ 19

3.4 Sound Intensity Measurements ....................... 19

4.2

4.3

4.4

Evaluation of Treatment Effectiveness 29

4.1 Description of Treatments .......................... 29

4.1.1 Torque Box (Structural Path) .................... 29

4.1.2 Additional Frames (Structural Path) ................ 30

4.1.3 Frame Damping (Structural Path) ................. 30

4.1.4 Pressure Bulkhead Double Wall (Aft Section Path) ........ 30

4.1.5 Sonic Fatigue Damping (Structural, Cabin Sidewall, and Aft Sec-

tion Paths) .............................. 30

4.1.6 Cabin and Cargo Skin Damping (Structural and Cabin Sidewall

Paths) ................................. 31

4.1.7 Floor Isolation (Structural Path) .................. 31

4.1.8 Aft Section Absorption (Aft Section Path) ............. 31

4.1.9 Torque Box Damping (Structural Path) .............. 31

4.1.10 Engine Dynamic Absorbers (Structural Path) ........... 31

4.1.11 Trim Panel Damping (Cabin Sidewall Path) ............ 31

Selection of Test Points ........................... 32

Acoustic and Vibratory Loads ........................ 32

4.3.1 Exterior Noise Levels ......................... 32

4.3.2 Pylon Vibration Levels ........................ 33

Cabin Noise and Vibration Levels ...................... 33

Analysis of Quiet Cabin Noise Levels 57

5.1 Acoustic and Vibratory Loads ........................ 57


iv

7

8

9

A

B


5.2 Cabin Noise Levels .............................. 58

Coml)arison of lOx8 and 8x8 Data 73

_;.I A,,,.._ti, _,.I Vil,v,.t.,,vy I,,m,l._ ........................ 73


6.1.2 Pylon and Fuselage Vibration Levels ................ 74


Comparison of Cruise and Non-Cruise Data 89

7.1 Acoustic and Vibratory Loads ........................ 89




Identification of Major Transmission Paths

8.1

8.2

8.3

8.4

8.5

109

Results of Sound Intensity Survey ..................... 109

Cabin Sidewall Path ............................. 110

Aft Section Path ............................... 111

Structural Path ................................ 113

Results of Partial Coherence Analyses ................... 114

8.5.1 Partial Coherence Analysis Concepts ................ 115

8.5.2 Analysis of 8x8 Data ......................... 116

8.5.3 Analysis of 10x8 Data ........................ 119

Summary and Conclusions 133

Estimation of The Turbulent Boundary Layer Pressure Wavenumber-

Frequency Spectrum 136

A.1 Definition of Wavenumber-Frequency Spectra ............... 136

A.2 Spectral Modeling of TBL Pressure Fluctuations ............. 138

A.3 Estimation Procedure ............................ 139

A.4 Flight Test Results .............................. 142

A.5 References ................................... 144

Partial Coherence Techniques 171

B.1 Mathematical Formulation .......................... 171

B.2 Flight Test Applications ........................... 174

List of Tables

2-1. UDF Engine Characteristics .................................................. 8

2-2. Test Point Smnmary ......................................................... 9

4-1. UHB Demonstrator Interior Treatment Configurations ....................... 35

4-2. Test Point Summary for Interior Treatment Evaluation ...................... 36

7-1. Test Point Summary for Non-Cruise Comparisons ........................... 93

8-1.

8-2.

8-3.

8-4.

Coherence Between Inputs and Outputs for BPF Tone (168 Hz),

Test Point 19550A03 ....................................................... 121

Reduction in BPF Tone Level at Interior Output Microphones

Conditioned with Exterior Microphone and Accelerometer Inputs,

Test Point 19550A03 ....................................................... 122

Coherence Between Inputs and Outputs for BPF Tone (168 and 210 Hz),

Test Point 19640B03 ....................................................... 123

Reduction in BPF Tone Level at Interior Output Microphones

Conditioned with Exterior Microphone and Accelerometer Inputs,

Test Point 19640B03 ....................................................... 124

A-1. Flight Test Conditions and Parameters .................................... 147

List of Figures

1-1. Demonstrator Noise Levels, Quiet Cabin Configuration ....................... 5

2-1. The MD-80 Aircraft ......................................................... 10

2-2. The UHB Engine and Pylon Installation ..................................... 11

2-3. The Demonstrator in Flight ................................................. 12

2-4. The Demonstrator on the Ground Showing Counter-Rotating Blades ......... 13

2-5. View from Under the Demonstrator Showing Relative Positions of the

JTSD and UDF Engines and Pylons ......................................... 14

2-6. The Furnished Cabin Showing Selected Microphone Locations ............... 15

2-7. UDF Propulsor Design Schematic ........................................... 16

2-8. Close Up View of UDF 10x8 Engine on the Demonstrator .................... 17

3-1. Exterior Microphone Locations on Fuselage and Pylon ....................... 21

vi

3-2.

3-3.

3-4.

3-5.

3-6,

3-7.

3-8.

Cabin Microphone Locations ................................................ 22

Aft Section Microphone Locations ........................................... 23

Location of Pylon and Fuselage Accelerometers .............................. 24

Location of Station 1271, Station 1322 and Aft Bulkhead

Accelerometers .............................................................. 25

A Typical Interior Cabin Seat Microphone Spectrum ........................ 26

Sound Intensity Survey Areas ............................................... 27

Location of Survey Grids .................................................... 28

4-1. Predicted Noise Transmission Paths for the MD-UHB Demonstrator ......... 37

4-2. Torque Box Design .......................................................... 38

4-3. Frame Modifications and Additions .......................................... 39

4-4. Pressure Bulkhead Double Wall Design ...................................... 40

4-5. Distribution of Sonic Fatigue Treatment Between Stations

1174 and 1510 .............................................................. 41

4-6. Engine Dynamic Absorber Locations on Mount Beam ........................ 42

4-7. Exterior Acoustic Loads Distribution of the 8X8 Propeller Blade Passage

Frequency for Each Test Point .............................................. 43

4-8. Fuselage Acoustic Loads: A-Weighted Broadband Levels ..................... 44

4-9. Acceleration Levels at BPF, 2BPF and UN1 Measured at the Inboard

Location of the Pylon Spar .................................................. 45

4-10. Maximum A-Weighted Levels at Seated Positions of Row 6 ................. 46

4-11. Maximunl A-Weighted Tone Levels at Seated Positions of Row 6 ............ 47

4-12a. A-Weighted Levels Measured at Microphone 6 in Row 6

(Window Seat, JT8D Side) .................................................. 48

4-12b. A-Weighted Levels Measured at Microphone 4 in Row 6

(Aisle Seat, JT8D Side) ..................................................... 49

4-12c. A-Weighted Levels Measured at Microphone 2 in Row 6

(Aisle Seat, UDF Side) ...................................................... 50

4-12d. A-Weighted Levels Measured at Microphone 1 in Row 6

(Window Seat, UDF Side) .................................................. 51

4-13. Measured Interior Noise Spectra in Row 6 for the Seat with

the Maximum A-Level (Configurations 1, 2 and 3) ........................... 52

4-14. Measured Interior Noise Spectra in Row 6 for the Seat with

the Maximum A-Level (Configurations 4, 5, 6 and 7) ........................ 53

4-15a. Frame Acceleration at BPF, 2BPF and UN1 at Fuselage

Station 1271, Longeron 2 .................................................... 54

4-15b. Frame Acceleration at BPF, 2BPF and UN1 at Fuselage


4-15c. Frame Acceleration at BPF, 2BPF and UN1 at Fuselage


vii

5-1. Variation of BPF Sound Pressure Level with Propeller Speed ................ 60

5-2a. Aft Engine Mount Accelerations at BPF, 2BPF and UN1 ................... 61

5-2b. Inboard Pylon Spar Accelerations at BPF, 2BPF and UN1 ................. 62

:, 3. A W,'ightcd lnteri,,r Nuisc Levels at R,,w 4 (Contiguration 4} ................ 03

5-4. A-Weighted Interior Noise Levels at Row 6 (Configuration 4) ................ 64

5-5. Maximum Measured Sound Pressure Levels at BPF, 2BPF, UN1

and JN1 Frequencies ........................................................ 65

5-6. Variation of the BPF A-Weighted Level in Row 4 and Row 6 ................ 66

5-7. Variation of the 2BPF A-Weighted Level in Row 4 and Row 6 ............... 67

5-8. Variation of the UN1 A-Weighted Level in Row 4 and Row 6 ................. 68

5-9. Variation of the JN1 A-Weighted Level in Row 4 and Row 6 ................. 69

5-10. Variation of the Tone Summations in Row 4 and Row 6 .................... 70

5-11. Variation of the Broadband Noise Levels in Row 4 and Row 6 .............. 71

5-12. Variation of the Overall A-Weighted Levels in Row 4 and Row 6 ............ 72

6-1. Comparison of 10x8 and 8x8 Spectra at Exterior Microphone 4 .............. 76

6-2a. Comparison of 10x8 and 8x8 BPF Tones on the Aft Fuselage ............... 77

6-2b. Comparison of 10x8 and 8x8 BPF Tones on the Aft Fuselage ............... 78

6-3a. Comparison of 10x8 and 8x8 Engine Mount Accelerations ................... 79

6-3b. Comparison of 10x8 and 8x8 Pylon Accelerations ........................... 80

6-4. Comparison of 10x8 and 8x8 Fuselage Acceleration Levels .................... 81

6-5a. Comparison of 10x8 and 8x8 Noise Levels at Row 6 Seat 1 .................. 82

6-5b.

6-5c.

6-5d.

6-5e.

6-6a.

6-6b.

Comparison of

Comparison of

Comparison of

Comparison of

Comparison of

Comparison of

10x8 and 8x8 Noise Levels at Row 6 Seat 3 .................. 83




10x8 and 8x8 Spectra at Row 6 Seat 4 ...................... 87

10x8 and 8x8 Spectra at Row 4 Seat 3 ...................... 88

7-1. Variation of

7-2. Variation of

7-3. Variation of

7-4. Variation of

8x8 Exterior Sound Pressure Level at BPF ...................... 94

10x8 Exterior Sound Pressure Level at BPF ..................... 95

10x8 Longitudinal Pylon Acceleration ........................... 96

10x8 Vertical Pylon Acceleration ................................ 97

7-5. Maximum Measured Sound Pressure Level in Row 6 at each BPF ............ 98

7-6a. Variation of 8x8 Interior Noise Components with RPM at

Row 6 Seat 1 ................................................................ 99

7-6b. Variation of 8x8 Interior Noise Components with RPM at

Row 6 Seat 3 ............................................................... 100

7-6c. Variation of 8x8 Interior Noise Components with RPM at

Row 6 Seat 4 ............................................................... 101

°,.

rill

7-6d. Variation of 8x8 Interior Noise Components with RPM at

Row 6 Seat 5 ............................................................... 102

7-6e. Variation of 8x8 Interior Noise Components with RPM at

R,,,, ,_ 'q¢,,t ,i |0_t

7-7a. Variation of IUx8 Interior Noise t_t_mpl, nents with RPM at

Row 6 Seat 1 ............................................................... 104

7-7b. Variation of 10x8 Interior Noise Components with RPM at

Row 6 Seat 3 ............................................................... 105

7-7c. Variation of 10x8 Interior Noise Components with RPM at

Row 6 Seat 4 ............................................................... 106

7-7d. Variation of 10x8 Interior Noise Components with RPM at

Row 5 Seat 1 ............................................................... 107

7-7e. Variation of 10x8 Interior Noise Components with RPM at

Row 5 Seat 6 ............................................................... 108

8-1. Measured Sound Intensities for the 100, 160, 200 and 315 Hz

One-Third Octave Bands, 8x8 Flight 2026 .................................. 125

8-2. Measured Sound Power Levels for the 100, 160, 200 and 315 Hz

One-Third Octave Bands, 8x8 Flight 2026 .................................. 126

8-3a. Cabin Sidewall Transmission Path: Exterior Noise Spectra

Measured on the Aircraft Fuselage, Test Point 19550A03 ................... 127

8-3b. Cabin Sidewall Path: Interior Noise Spectra, Test Point 19550A03 ......... 128

8-4. Aft Section Transmission Path: Unpressurized Section Noise Spectra,

Test Point 19550A03 ....................................................... 129

8-5. Aft Section Transmission Path: Comparison of Interior Noise Spectra

at Row 6, Seat 1 ........................................................... 130

8-6. Comparison of Noise and Vibration Data Acquired in the Aft Section

of the Aircraft, Flight 1955 ................................................. 131

8-7. Transducers Used in the Partial Coherence Analysis of 8x8 Data ............ 132

A-1. Sketch of the Wavenumber-Frequency Spectrum ............................ 148

A-2a. Sound Pressure Level Spectra for the Probe Microphones

(Flight Condition 1) ....................................................... 149

A-2b. Sound Pressure Level Spectra for the Probe Microphones


A-2c. Sound Pressure Level Spectra for the Probe Microphones


A-2d. Sound Pressure Level Spectra for the Probe Microphones


ix

A-2e. Sound Pressure Level Spectra for the Probe Microphones


A-3a. Estimated Wavenumber-grequency Spectrum (Flight Condition 1) ........ 154

A-3b. Estimated Wavenumber-Frequency Spectrum (Flight Condition 2) ........ 155

A-3c. Estimated Wavenumber-Frequency Spectrum (Flight Condition 3) ......... 156

A-3d. Estimated Wavenumber-Frequency Spectrum (Flight Condition 4) ........ 157

A-3e. Estimated Wavenumber-Frequency Spectrum (Flight Condition 5) ......... 158

A-4a. Wavenumber-Frequency Contour Plot for TBL Pressure Spectrum

(Flt. Cond. 1) ............................................................. 159

A-4b. Wavenumber-lh'equency Contour Plot for TBL Pressure Spectrum

(Flt. Cond. 2) ............................................................. 160

A-4c. Wavenumber-Frequency Contour Plot for TBL Pressure Spectrum

(Fit. Cond. 3) ............................................................. 161

A-4d. Wavenumber-Frequency Contour Plot for TBL Pressure Spectrum

(Fit. Cond. 4) ............................................................. 162

A-4e. Wavenumber-Frequency Contour Plot for TBL Pressure Spectrum

(Fit. Cond. 5) ............................................................. 163

A-Sa. Wavenumber Variation of TBL Pressure Spectral Density (F=50 Hz) ...... 164

A-Sb. Wavenumber Variation of TBL Pressure Spectral Density (F=75 Hz) ...... 165

A-Sc. Wavenumber Variation of TBL Pressure Spectral Density (F=100 Hz) ..... 166

A-Sd. Wavenumber Variation of TBL Pressure Spectral Density (F=150 Hz) .... 167

A-6. Comparison of Flight Test TBL Pressure Data with Corcos' Model ......... 168

A-7. Comparison of TBL Pressure Spectral Density with Laganelli's Model ...... 169

A-8. Comparison of Flight Test TBL Pressure Data with Witting's Model ....... 170

B-1. Example of the Partial Coherence Conditioning Process for Output

Microphone 5, 8x8 Test Point 19550A03 .................................... 175

X

1 Introduction

1.1 Background

Over tile past fifteen years there has been a renewed interest in the use of fuel-

efficient turboprop engines for aircraft powerplants, due to the concern over rising

fuel costs and the accompanying desire for energy conservation. Advanced turboprop

engines, which make use of recent developments in blade design and fabrication, have

the potential for providing thrust that is comparable to today's turbofan engines with

a significant reduction in fuel use.

McDonnell Douglas Corporation (MDC) is currently developing a new generation

of commerical transport aircraft that would be powered by such advanced turboprop

engines, designated as Ultra High Bypass (UHB) engines. MDC's plans are to develop

a derivative of the MD-80 series aircraft, which would incorporate two aft-mounted

UHB engines in place of the two turbofan engines now utilized.

One of the major concerns associated with UHB aircraft is the noise environment

that may be experienced by passengers. The longest portion of a typical flight is spent

at high altitude, high speed cruise. Under these conditions the UHB engine is expected

to produce its highest noise levels since the tips of the propeller blades are moving

at supersonic speeds. The resulting acoustic energy is generated in discrete tones, at

frequencies corresponding to the blade passage frequency and its multiples for each

propeller rotor. These blade passage frequencies lie typically between 100 and 250

Hz, where the transmission loss characteristics of standard aircraft sidewalls are not

sufficient to reduce the high noise levels expected on the fuselage exterior to acceptablylow interior levels.

Ill 1985, a UHB Technology Readiness Program was initiated at Douglas Aircraft

Company to address the various technical issues associated with the development of

UHB aircraft, including interior noise. A major element of this program was a series of

demonstration flight tests using a modified MD-80 test aircraft with a prototype UHB

engine in place of the left JT8D turbofan engine. One of the primary purposes of the

flight tests of this MD-UHB Demonstrator aircraft was to prove that a "Quiet Cabin"

could be achieved in a commercial transport aircraft powered by advanced turboprop

engines. For these tests, a Quiet Cabin meant that the maximum noise level in the

cabin during high altitude, high speed cruise conditions would not exceed 82 dBA.

In order to meet this interior noise goal for the MD-UHB Demonstrator, the ap-proach adopted was to:

1. Estimate interior noise levels for an untreated aircraft, using projections of UHB

exterior noise levels and vibration loads;

2. Propose candidate noise control treatments applicable to the various expectedairborne and structureborne transmissionpaths which would reduceinterior noiselevelsto meet the goal of 82 dBA;

3. Evaluate the effectivenessof the candidate treatments using fuselagegroundtests; and

4. Progressively install and evaluate the most promising treatments on the Demon-

strator aircraft.

This approach also provided noise and vibration data in flight for various treatment

configurations, from which treatment effectiveness and noise transmission paths into

the cabin could be studied.

Starting in October 1986, a series of interior noise control tests was conducted in a

fuselage ground test facility to evaluate selected treatments planned for installation on

the MD-UHB Demonstrator. Flight tests of the Demonstrator powered by one JT8D

engine and one UHB engine began in June 1987, with a minimal noise control treatment

package. The UHB engine used for these tests was a General Electric "Unducted Fan"

(UDF) engine, which utilizes the exhaust from a small jet engine to drive two rows of

highly swept, counter-rotating propeller blades. For the first several flight tests, the

UDF engine configuration included eight blades on both the forward and aft rotors (an

8x8 configuration).

Additional noise control treatments were then added, aided by the ground test re-

suits. In July the installation of the full Quiet Cabin treatment package was completed,

and the interior noise goal of 82 dBA was attained. In August, the 8x8 UDF engine

configuration was replaced by a 10x8 configuration (10 blades on the forward rotor and

8 blades on the aft rotor), the 82 dBA interior noise goal was again achieved. Figure 1-1

illustrates the interior noise environment measured on the Demonstrator during one of

the 10x8 flights. The measurements showed the noise levels throughout the aft cabin

to be in a narrow range from 78 to 82 dBA; these levels are below the levels in the aft

cabin of many current jet transports. Furthermore, observers on the aircraft (including

representatives from the airlines and the press) noted that the propeller blade passage

tones were not perceptible in flight, thus alleviating the early concern that these tones

would render the cabin noise environment unacceptable.

In subsequent phases of the program the 8x8 configuration was reinstalled for ad-

ditional testing, and in February 1988 selected treatments were removed to investigate

the potential for reducing treatment weight while maintaining low cabin noise levels.

The entire MD-UHB Demonstrator flight test activity (including flights to study

aerodynamic and structural dynamics issues and for marketing purposes) totaled 93

flights for 165.5 hours, over a 10-month period ending in March 1988. The fight tests

2

proved that UHB engine technology could successfully be applied to commercial trans-

port aircraft. In particular, the achievement of a comfortable cabin noise environment

on a UHB-powered aircraft, comparable to that on a turbofan-powered aircraft, was

demonstrated to the technical community, the airlines, and the public.

As one element of NASA's Advanced Turboprop (ATP) technology program, the

NASA/Industry flight demonstration program was started in 1987 to work with the

aircraft industry in the study of passenger cabin noise in advanced turboprop aircraft.

Under this program, a contract with MDC required detailed analysis of the interior

noise measurement data acquired during the several ground and flight tests described

above. This report describes the results of MD-UHB Demonstrator flight tests. (The

results of the fuselage ground tests are documented in NASA CR 181819.)

1.2 Objectives and Approach

The objectives of the analyses described in this report are to investigate the interior

noise characteristics of advanced turboprop aircraft with aft-mounted engines and to

study the effectiveness of selected noise control treatments in reducing passenger cabinnoise.

These objectives were accomplished by analyzing the noise and vibration data col-

lected during numerous MD-UHB Demonstrator flight tests. The majority of the data

were obtained at high altitude, high speed cruise conditions with the 8x8 engine con-

figuration. Exterior and interior noise level distributions were determined from these

data, for several noise treatment installations. Additional data were obtained under

comparable cruise conditions with the 10x8 engine configuration, and for both configu-

rations at lower speeds and altitudes. Differences in levels for these alternate conditions

and configurations were determined relative to the 8x8 cruise data.

The relative strengths of airborne and structureborne transmission paths were de-

fined using the results of a sound intensity survey, combined with measured cabin noise

levels and exterior noise and vibration data measured on the fuselage and pylon. Partial

coherence analysis techniques were also used to provide further insight into the noise

transmission paths.

Since broadband noise due to the turbulent boundary layer was found to be a sig-

nificant contributor to the total interior noise level, a supplementary study of this

noise source was undertaken. Estimates of the turbulent boundary layer pressure

wavenumber-frequency spectrum were made with a new spectral analysis technique

using noise levels measured on the fuselage surface, and compared with theoretical

models and empirical models based on wind tunnel data.

3

1.3 Report Overview

This report documents the measurement and analysis procedures for MD-UHB

Demonstrator noise and vibration flight test data related to passenger cabin noise,

and describes the test results. The next two sections of this report discuss the flight

test program and data acquisition and analysis, respectively. In Section 4, the various

noise control treatments installed on the Demonstrator are defined and their effective-

ness is evaluated. Section 5 presents the results of measurements in the fully treated

"Quiet Cabin", for high speed, high altitude cruise conditions with the 8x8 engine

configuration. Sections 6 and 7 compare measurement results for the 10x8 engine con-

figuration and for non-cruise conditions, respectively, with the Section 5 results. The

following section identifies the major transmission paths (airborne and structureborne)

into the cabin. The final section summarizes the conclusions of this study. In Appendix

A, results of the wavenumber-frequency analysis of turbulent boundary layer noise are

presented. Appendix B provides the mathematical background for the partial coher-

ence analysis used in Section 8. Tables, figures, and references are located at the end

of each section or Appendix.

4

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0

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0

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2 Flight Test Program

2.1 Test Aircraft

The Demonstrator aircraft is a modified MD-80 series aircraft. The MD-80 (see

Figure 2-1), characterized by an overall length of 147.8 ft and a wingspan of 107.8 ft,

is designed to carry 155 passengers on short-to-medium range routes. The aircraft is

normally powered by two Pratt & Whitney JT8D-209 engines. The JT8D-209 is a

dual-rotor, fully ducted turbofan with a low bypass ratio of 1.8.

The Douglas Aircraft-owned MD-80 (ship 909) was modified by replacing the left

engine with a General Electric proof-of-concept Unducted Fan (UDF). Although the

nacelles of the JT8D-209 and the UDF are of comparable diameter, the large diameter

of the propellers required that the thrust axis be moved outboard and higher than that

of the JTSD-209 to provide sufficient blade tip-to-fuselage and tip-to-ground clearance.

Accordingly a new pylon was installed with an approximate 5 ft span to replace the 9

inch stub pylon used with the JTSD engine, and stonger fuselage frames were installed

at the forward and aft pylon spar locations to support the new structure. The relative

positioning of the JT8D and of the UDF is illustrated in Figure 2-2. Figures 2-3, 2-4,

and 2-5 are photographs showing different views of the Demonstrator aircraft.

The last 20 ft of the cabin was configured with a current generation MD-80 interior,

including seats, baggage racks, interior lighting and trim, and a functional lavatory.

This was done to provide the most realistic environment for noise measurements and

to provide an attractive and comfortable interior environment for demonstration flights.

Flight test instrumentation was located forward of this cabin area. Figure 2-6 shows

the furnished aft cabin, with microphones mounted at selected seat and aisle locations.

2.2 Propulsion System

The UDF engine consists of an F404 low bypass ratio turbofan thermodynamically

coupled to the aft mounted propellers. Figure 2-7 illustrates how the core engine

exhausts are driven through a set of counter-rotating turbines at the base of the two

counter-rotating propellers. Figure 2-8 is a close-up view of the engine mounted on the

Demonstrator.

During the flight test program, two propeller configurations were used: eight-bladed

propellers on both the forward and aft rotors (designated as an 8x8 configuration), and

a combination ten-bladed propeller on the forward rotor and eight-bladed propeller on

the aft rotor (designated as a 10x8 configuration). The pertinent characteristics of the

UDF engine with both rotor configurations are summarized in Table 2-1.

6

2.3 Interior Noise Test Points

Of the 93 flights in the test program, several were primarily oriented to interior noise

measurenlent purposes. Interior noise data were also acquired on several flights that

were oriented to other purposes. Table 2-2 lists all the flights and test points in which

interior data were collected. As this table shows, data were obtained for both engine

configurations, for seven interior configurations (defined in Table 4-1 and described in

detail in Section 4), and several altitude, Mach No., and engine speed combinations.

UDF

Table 2-1

Engine Characteristics

10x8 Configuration

Forward Fan Diameter:

Aft Fan Diameter:

Forward Fan:

Aft Fan Rotation:

Number of Blades:

3.556m(11.67 ft.)

3.353m(11.00 ft.)

CCW looking forward

CW looking forward

10 fwd, 8 aft

8x8 Configuration

Forward Fan Diameter:

Aft Fan Diameter:

Forward Fan:

Aft Fan Rotation:

Number of Blades:

3.56m(11.67 ft.)

3.25m(10.67 ft.)

CCW looking forward

CW looking forward

8 fwd, 8 aft

Test

Table 2-2

Point Summary

DATE UHB RUN FUSELAGE ALTITUDE MACH

CONFIG. NUMBER CONFIG. (ft.) NUMBER

6-25-87

6-26-87

6-27-87

7-7-87

8-26-87

8-26-87

2-22-88

3-05-88

3-19-88

3-21-88

3-22-88

8x8

9'

8x8

8x8

.t

.1

8x8

10x8

10x8

8x8

8x8

8x8

8x8

8x8

19510001

19510002

19510003

19510004

19520A01

19520A02

19520A03

19520A04

19520A05

19530A01

19530A02

19530A03

19530A04

19530A05

19530A06

19530A07

19530A08

19550A01

19550A02

19550A03

19550A04

19550A05

19550A06

19550A07

19640A01

19640A02

19640A03

19640A04

19640A05

19640A06

19640A15

19640B01

19640B02

19640B03

19640B04

2O160HO1

20160H02

20160H05

20160H06

2026*

2O33OKOI

2o33OKO2

20330K03

2O330K04 i

20340K01

20340KO2

20340K03

20340K04

20350001

ROTOR SPEED

(RPM)

1280

1265

1260

1250

1270

1270

1260

1250

1270/1170

1270

1260

1250

1225

1270

1260

1250

1230

1285

1270

1260

1250

1225

1200

1170

1200

1210

1230

1250 "

1265 "

1290

1305

1275 4

1265 "

1255 "

1240 "'

1250

1255 "

1245 ""

1245 ""

1200 5

1255 5

1255 ""

1255 ""

1350 "

1255 6

1265

1285

1285

1265 7

Sound Intensity Survey Flight

35000

35000

35000.4

11

ti

35000

22500

.1

31000

35000

0.76

0.76

0.76

0.77

0.76

0.76

0.55

0.71

0.76

0.75

5 35000 0.76

19 0.77

" 0.76

33500 0.76

22500 0.69

" 0.70

0.77

35000 0.76

0.81

1' 0.84

"" 0.86

t 35000 0.77

9

FIGURE 2-1. The MD-80 Aircraft

10

; " !

\oo

11

• % • 7

{; }:

12ORIGINAL PAGE

BI..AC_,,KAND WHITE PHOTOGRAPH

]3 ORIGINALPAGEBLACK AND WHITE PHOTOGRAPt4

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=

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FIGURE 2-5. View from Under the Demonstrator Showing Relative Positions

of the JT8D and UDF Engines and Pylons.

ORIGINAL _";_,GE 14

BLACK AND V,/HITE PHOTOGRAPH

15ORIGINAL PAGE

BLACK AND WHITE PHOTOGRAPH

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17 ORIGINAl PAGEBLACK AND WHITE PHOTOGRAPFf

3 Data Acquisition and Analysis

3.1 Acquisition System

The Demonstrator noise and vibration data were acquired with an airborne digital

system capable of recording simultaneously up to 256 channels of digital time series

data at a rate of 6400 samples per second, with a dynamic range of 80 dB. Anti-aliasing

filters were chosen to provide a useful frequency band of 0 to 2000 Hz. The data were

typically acquired over a period of 15 seconds.

3.2 Transducers

Since the overall objectives of the Demonstrator test program addressed many

concerns including acoustics, stress, sonic fatigue, and vibration, transducers installed

throughout the aircraft consisted of numerous microphones, accelerometers, and strain

gauges. Of these, transducers useful for acquiring data for interior noise analysis

purposes included selected exterior and interior microphones, and frame, pylon, and

bulkhead accelerometers. Microphone and accelerometer measurement locations often

changed during the test program, to accomodate the requirements of a specific flight.

In the following, the location of all transducers used throughout the test program for

interior noise-related measurements are described; transducer locations used for specific

test points are defined in subsequent sections.

Exterior microphones on the fuselage surface were located in the forward and aft

propeller planes and at several additional points on the fuselage surface. Microphones

were also located on the upper and lower surfaces of the UHB pylon. Figure 3-1 shows

these exterior microphone locations.

Interior microphones were located in the aft cabin and in the unpressurized sec-

tion aft of the pressure bulkhead. The cabin microphones were located at positions

approximating that of both seated and standing passengers, as shown in Figure 3-2.

In each row, microphones at seat positions were 40 inches above the floor, and at aisle

positions were 72 inches above the floor on the aircraft centerline. The locations of the

microphones in the unpressurized aft section are shown in Figure 3-3.

Accelerometers were located at the foward and aft pylon spars and engine mount

points, and on the frame at station 1437 in the forward propeller plane (see Figure

3-4). Accelerometers were also located at several points on the frames at stations 1271

and 1322, and on the aft pressure bulkhead; these locations are shown in Figure 3-5.

18

3.3 Data Processing and Analysis

Processing of tile test data was conducted interactively with specially developed

software. Frequency spectra with a bandwidth of 3.125 Hz (or 0.78125 Hz when required

for finer resolution) were generated by taking the autospectrum of the time series data.

Ttle measured spectra are typically characterized by several tones superimposed on

a broadband background, as shown in the sample interior noise spectrum in Figure

3-6. The frequencies of the observed tones are related to the shaft rotational speeds of

the two engines and to the blade passage frequencies associated with the two propeller

rotors, plus harmonics of these various frequencies.

The rotating lnachinery ill the engines of the Demonstrator aircraft generate tones

at several frequencies. The Pratt & Whitney JT8D-209 is a dual-rotor turbofan and

generates tones at two primary frequencies. The first is associated with the rotational

speed (N1) of the low pressure stage and the second with the rotational speed (N2) of

the high pressure stage of the engine. In this report these two frequencies are designated

as JN1 and JN2, respectively. Similarly, tones are generated at frequencies associated

with the rotational speeds of the low and high pressure stages of the UDF engine, and

herein are designated as UN1 and UN2, respectively. In addition, tones are generated at

the frequency associated with the rotational speed of the rotor shaft of each propeller.

This frequency is usually referred to as the lp frequency of the propeller; in this report

the shaft speeds of the two rotors are nominally the same, so that only one lp tone

Occurs.

_Ikmes at tile blade passage frequency (BPF) and harnmnics are generated by each

propeller. F_,r the 8x8 engine configuration, each eight-bladed rotor generates a tone at

a BPF corresponding to a frequency eight times that of tile lp frequency. Harmonics

are labelled as 2BPF, 3BPF, etc. Tile spectrum shown in Figure 3-6 clearly exhibits the

BPF and 2BPF tones. For the 10x8 engine configuration the forward ten-bladed rotor

generates a tone at a BPF corresponding to ten times that of the lp frequency, while

the BPF tone from the aft eight-bladed rotor is at the same frequency (eight times lp)

as for the 8x8 configuration. To distinguish the two 10x8 blade passage frequencies, in

this report they are designated as BPF(10) and BPF(8).

3.4 Sound Intensity Measurements

In addition to the digital noise and vibration data acquired routinely throughout

the flight tests, a sound intensity survey was conducted during a dedicated flight (flight

2026).

For this test, the cabin sidewalls, ceiling, floor, right and left engine pyhm bulkheads

19

and aft pressure bulkhead were divided into survey areas. Figure 3-7 shows the extent

of the survey, and Figure 3-8 defines the actual grid locations on each surface. The

typical grid area is approximately 7 to 8 square feet.

The sound intensity survey was conducted using Norwegian Electronics Model 216

Sound Intensity Probes. These probes utilize a single microphone to measure sound

pressure, and an ultrasonic transducer to measure particle velocity. Each area was

scanned approximately four inches away from the surveyed surface. Norwegian Elec-

tronics Model 830 Real Time Analyzers collected and processed the data to yield one-

third octave band sound intensity levels, from which one-third octave band sound power

levels were computed.

2O

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FIGURE 3-3. Aft Section Microphone Locations.

23

./.......

FUSELAGEACCELEROMETER

X LATERAL

y LONGITUDINAL

Z VERTICAL

ACCELEROMETER DESCRIPTION

ACCEL# LOCATION DIRECTION

1 aftfuselage radial

2 fwd. pylon spar vertical

3 aftpylon spar vertical

4 fwd. enginemount vertical5-X lateral

5-Y aftenginemount longitudinal

5-Z vertical

FIGURE 3-4. Location of Pylon and Fuselage Accelerometers.

24

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co')

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0

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UpperSidewall

Bag Ra,

UpperSidewall

Lower

Sidewall

Lower

Sidewall

Floor

LOOKING AFT

|-Pylon Pylon [11

FIGURE 3-7. Sound Intensity Survey Areas.

27

AftBulkheads

Sb 1303-

Sh 1269o

Sb I_I-

on Att B'emureButkh_

LefL

Sidewall

Sb I1_ -

Sb II_ -

Sto1174-

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Aft

Floor'

FIGURE 3-8. Location of Survey Grids.

28

4 Evaluation of Treatment Effectiveness

Ill order to achieve the UHB Demonstrator's interior noise goal of 82 dBA or less at

any seal. location during nornlal cruise conditions, several noise reduction treatments

were designed and installed on the aircraft. These treatments were selected to con-

trol noise propagation along three potential paths (see Figure 4-1): an airborne path

through the cabin sidewall, an airborne path into the unpressurized aft section and

then through the pressure bulkhead, and a structural path through the engine pylon

and into the fuselage structure. All treatments complied with FAA FAR 25 regulations

and did not alter the standard dimensions and layout of the cabin.

The original plan for treatment evaluation was to install selected treatments or

sets of treatments in phases, so that the effectiveness of each installation could be

measured. However, this approach was constrained by the extensive time required

for aircraft modifications associated with several of the treatments, and the limited

number of flights available for interior noise tests. As a result, a number of treatments

were installed at the same time, and comprised the "Baseline" configuration (i.e., the

cabin configuration for the first interior noise test flight). Additional treatments were

subsequently added, up to the fully configured Quiet Cabin, when the goal level of 82

dBA was achieved. Selected treatments beheved to have httle effect on cabin noise

levels (based on prior flight and ground tests) were removed for later flights in an

attempt to maintain the low noise level while reducing treatment weight.

In this section, treatment effectiveness is evaluated based on analysis of changes in

cabin noise levels after each treatment is installed. In section 5, a detailed analysis of

exterior and interior levels measured for the Quiet Cabin configuration is presented.

Some of the Quiet Cabin results are used in this section to identify the major test

l)arameters to be studied in this evaluation.

4.1 Description of Treatments

Table 4-1 lists the flight tests during which interior noise data were collected for the

8x8 UDF engine configuration, along with the interior noise treatment configuration

number and a list of the treatments included. Each treatment is described below.

4.1.1 Torque Box (Structural Path)

The torque box is a device designed to greatly increase the stiffness of an existing

frame, particularly to torsional motion. It consists of a new frame installed approxi-

mately four inches from an existing frame (located at station 1287), with a cover plate

29

over both frames (see Figure 4-2).

4.1.2 Additional Frames (Structural Path)

Four additional frames, identical to current production frames, were installed mid-

way between existing frames in the aft cabin area. Figure 4-3 shows the locations and

station numbers. The purpose of these frames is to provide additional sidewall stiffness.

4.1.3 Frame Damping (Structural Path)

Damping material was applied to the new frames, and to several existing frames to

reduce structural vibration. Figure 4-3 illustrates the method of installation and the

locations of the treated frames.

4.1.4 Pressure Bulkhead Double Wall (Aft Section Path)

To reduce sound transmission from the aft unpressurized fuselage section into the

cabin, a double wall was constructed of 0.063 inch aluminum, and installed approx-

imately three inches forward of the aft pressure bulkhead. The configuration also

includes damping material applied to the pressure bulkhead, and isolator mounts to

attach the double wall to the fuselage structure. See Figure 4-4 for details.

4.1.5 Sonic Fatigue Damping (Structural, Cabin Sidewall, and Aft Section

Paths)

A damping treatment was used to reduce sonic fatigue of the fuselage skin in the

vicinity of the propeller planes where high acoustic loads were expected. Since this

treatment affects noise transmission into the aft section and structureborne noise prop-

agation into the cabin, it is listed as a noise treatment. The damping material was

installed on the skin between the frames and longerons, starting in the aft section at

station 1510, through the passenger cabin to station 1174. The treatment is distributed

on the skin from the bottom of the fuselage, extending across the UDF side, and ending

at a point 30 ° from the top of the aircraft on the 3T8D side (see Figure 4-5).

3O

4.1.6 Cabin and Cargo Skin Damping (Structural and Cabin Sidewall

Paths)

A second damping treatment was applied on all skin panels throughout the pas-

senger cabin and the cargo compartment from stations 1098 to 1338. The treatment

extended around the entire circumference of the fuselage. This treatment was added

over the areas where sonic fatigue damping material was already appfied.

4.1.7 Floor Isolation (Structural Path)

Floor panels from stations 1090 to 1350 were installed on isolation mounts. The

isolator damping material was the same as as used on the fuselage frames.

4.1.8 Aft Section Absorption (Aft Section Path)

To absorb low frequency noise in tile aft, unpressurized section, 2.0 inch thick quilted

fiberglass insulation blankets were installed over the pressure bulkhead and over the

ventral stair cage.

4.1.9 Torque Box Damping (Structural Path)

Tests of the effectiveness of the torque box in Douglas' fuselage ground test facility

indicated that this treatment could increase, rather than decrease, cabin noise levels.

In an attempt to correct this situation, a damping treatment was applied to the cover

plate of the torque box. The installation was similar to the damping installation used

o11 the additional frames (see Section 4.1.3 above).

4.1.10 Engine Dynamic Absorbers (Structural Path)

Three tuned absorl)ers (see Figure 4-6) were added to the rear engine mounts to

reduce structureborne noise, and in particular pylon vibration transmissibility. The

absorbers were tuned to the nominal cruise operation blade passage frequency (168

Hz).

4.1.11 Trim Panel Damping (Cabin Sidewall Path)

From station 1098 to station 1303, damping material was applied to the trim panels

within the sidewall cavity, in order to reduce panel resonant response and increase mass

31

density.

4.2 Selection of Test Points

To evaluate the effectiveness of the interior noise treatments, data acquired during

several flights were compared. A total of seven configurations were investigated during

seven different flights. To ensure the validity of this analysis, test data were selected to

correspond to flight conditions with nearly identical characteristics (pressure altitude,

airspeed and rotor rpm). Table 4-2 summarizes the available test points and highlights

those chosen for this study.

4.3

4.3.1

Acoustic and Vibratory Loads

Exterior Noise Levels

Analysis of the Quiet Cabin data shows, among other characteristics, that :

• The propeller blade passage frequency (BPF) is the dominant propeller noise

contributor to the interior noise.

• Exterior BPF levels are highest in the two prop planes, then decrease rapidly

with distance forward of the prop planes.

• The boundary layer noise is a significant contributor to the interior noise, and is

transmitted through the fuselage sidewall.

To accurately study treatment effectiveness using interior noise data acquired during

multiple flights, it is therefore important to verify that exterior levels for the BPF tones

and the boundary layer are essentially constant from flight to flight.

Figure 4-7 presents BPF tone levels measured on the fuselage surface. It can be

observed that in the prop planes, the variability of the BPF tone level among the

several flights is small (2 to 3 dB). The distribution of BPF levels along the fuselage

is very consistent from flight to flight, showing acceptably small variations in level at

each location.

Concerning boundary layer noise, Figure 4-8 presents the A-weighted broadband

noise levels measured at various exterior locations along the cabin sidewall for each

configuration. Again the measured levels are nearly identical from flight to flight.

32

4.3.2 Pylon Vibration Levels

Vibration data measured on the pylon spars show the presence of strong tones at

the UN1 frequency as well as at the propeller BPF and 2BPF. Since a structureborne

path along the pylon could be a contributor to the interior noise levels for these tones,

it is important to also verify the consistency of the vibration data for all flights.

Figure 4-9 presents the BPF, 2BPF, and UN1 vibration levels for each configuration

as measured on the inboard side of the forward pylon spar and aft pylon spar. It is

readily observed that the vibration amphtudes did not vary significantly from flight to

flight.

4.4 Cabin Noise and Vibration Levels

Noise levels were monitored in the cabin in seat row 6 for all seven test points, per-

mitring a comparison of levels at identical locations for all treatment configurations.

(Measurements were also obtained in either seat row 4 or seat row 5, depending on

flight; these data are not included below because of the inconsistent locations.) Figure

4-10 presents the maximum overall A-weighted levels measured among the five seat lo-

cations in seat row 6 for each configuration, and the corresponding tone and broadband

component levels. (The tone component is determined by summation of the various

contributing tones; the broadband component is determined by subtracting the tone

component from the overall.) The following observations can be drawn from this figure:

• Treatment configurations 2 (torque box damping) and 3 (engine dynamic ab-

sorbers) have a negligible effect o11 the nlaximum interior noise level,

• Tre,'ttment contiguration 4 (cabin filrnishings with damped triln panels) reduces

that level significantly, and

• Treatment configurations 5, 6, and 7 (removal of treatments previous to configu-

ration 4) have negligible effect on interior noise levels.

Figure 4-11 similarly presents the maximum A-weighted tone levels in seat row six

observed for each configuration, at the BPF, 2BPF, UN1, and JN1 frequencies (all of

which contribute to interior noise levels.) For the BPF, 2BPF, and JN1 tones, the same

general observations can be made from the figure as from Figure 4-10, except for the

increase in 2BPF levels for configurations 6 and 7 (which may indicate the contribution

of the aft section path for this frequency when treatment is removed; see section 8.2).

For the UN1 tone, the variation in level does not follow the same trends (but rather

varies with nearby frame acceleration, see below).

33

The noise variation with cabin configuration is further investigated in Figures 4-

12a, b, c, and d for selected seat locations in row 6. Again, the A-weighted overall

sound levels as well as tone and broadband components are presented. An immediately

recognizable trend is the noise reduction caused by the addition of the trim panels

and furnishings (configuration 4), and the typically negligible change in level for the

other configurations. The tone summation at the window seat on the UDF side (Figure

4-12d) does not however conform to this trend and exhibits only small changes in level

among all configurations; this behavior may be the result of the modal characteristics

of the cabin. Although the relative levels of the tone and broadband components vary

from seat to seat, it is noted that the broadband contribution is always equal to or

greater than the tone summation.

The interior noise spectra corresponding to the data in Figure 4-10 are presented

in Figures 4-13 and 4-14. Figure 4-13 presents the spectra for configurations 1 through

3. The spectra for the Quiet Cabin and successive configurations (4 through 7) are

shown in Figure 4-14. In Figure 4-13, the BPF tone levels at 168 Hz for all three

configurations are between 90 and 95 dB. Figure 4-14 data show BPF levels between 85

and 90 dB. Similarly the broadband levels were reduced by 5 to 10 dB after installation

of configuration 4. In general the spectral characteristics are similar among the three

spectra in Figure 4-13; among the four spectra in Figure 4-14 the spectral characteristics

are also generally similar.

Figures 4-15a, b, and c are plots of acceleration versus configuration for selected

accelerometers placed on frame 1271. BPF, 2BPF, and UN1 tone levels are presented as

they are the predominant tones observed on this frame. The BPF tone levels appear to

vary randomly with configuration while the 2BPF tone levels remain relatively flat. No

specific trends with treatment configuration can be identified for these two frequencies.

For UN1 tones, a possible effect of the engine dynamic absorbers on the acceleration

levels can be seen (levels generally decrease when absorbers are installed and increase

when absorbers are removed at two of the three measurement locations). However, since

the absorbers are tuned to 168 Hz and UN1 is 185 Hz, it is likely that the observed

variation in acceleration level is unrelated to this treatment.

The acceleration of this frame at BPF and 2BPF appears unrelated to cabin noise

levels, however the frame acceleration and cabin noise levels at UN1 show similar

variations with configuration (see Figures 4-11 and 4-15a, c).

34

TABLE 4-1 UHB Demonstrator Interior Treatment Configurations.

Flight No. Configuration No.1951

1952

1953

1955

2016

2034

2035

1

(Baseline)

4

(Quiet Cabin)

5

6

'"7

Treatment Description

Torque BoxAdditional Frames

Frame DampingDouble Wall Bulkhead

Sonic Fatigue Damping

Cabin Skin Damping

Cargo Skin DampingFloor Isolation

Aft Section Absorption

(Cabin Unfurnished, with sidewallthermal insulation but without

trim panels)# 1 plus:

Torque Box Damping# 2 plus:

Engine Dynamic Absorbers3 plus:

Cabin FurnishingsTrim Panels

Trim Panel Damping

# 4 less:

Engine Dynamic Absorbers

# 5 less:

Aft Section Absorption

Cargo Skin Damping

# 6 less:Double Wall Bulkhead

(section over pressurebulkhead door)

35

TABLE 4-2

Test Point Summary for Interior Treatment Evaluation.

Rul,tone,hiNumber Num.

19510001 1

2

3

4

19520A01 2

2

3

4

(spSt)519530A01 3

2

3

4

5

67

8

19550A01 4

23

4

56

7

20160H01 5

2

5

6

20340K01 6

20350001 7

N3/N4 Rotor Speed (rpm)

1170 I 120011225 1 1245 [ 12501 1255 [ 1260 I, 1265 ] 1270 ] 1280

®

®

®

®

®

®

• Available Test Points

®

@ Test Points Selected for Treatment Effectiveness Study

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FRAME CHANNELS

DAMPING MATERIAL - ON SKIN BETWEEN LONGERONS

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FIGURE 4-2. Torque Box Design.

38

SIDEWALL STIFFNESS/DAMPING MOD

ENGINE MOUNT

NEW FRAME BULKHEAD

LEFT SlOE

LOOKING INBOARD

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MATERIAL

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MATERIAL

FIGURE 4-3. Frame Modifications and Additions.

39

DOUBLE WALL PRESSURE BULKHEAD

ATTACHED TO SKIN AND PRESSURE BULKHEAD

WITH ISOLATION MOUNTS

PRESSURE DOME

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FIGURE 4-4. Pressure Bulkhead Double Wall Design.

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51

C)cr_

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BPF

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2BPF

I I J i

Configuration 1

ORSPL - 97.9 dB

RLEVEL - 86.1 dB[R]

II

i

i q

.II li_ J

L I

• ! ' YI -V_| I

Configuration 2

OASPL - 98.8 dB

ALEVCL - 86.2 dBtRJ

I

1.... I .... I .... I .... I .... 1.... I .... I .... I .... I ..... I0 50 t O0 350 _lO0 450 5(10

Config_ration 3

ORSPL - 98.0 dB

RLEVCL - 86.4 d_lAl

[50 200 250 300

Frequerm,4. (Hz]

FIGURE 4-13. Measured InteriorNoise Spectra in Row 6 for the Seat

with the Maximum A-Level (Configurations I, 2 and 3).

52

o

o

(:3

CD

o.

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CD.

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t .... t .... I .... t .... I .... I .... I .... I .... I .... I .... I0 50 100 150 200 250 300 350 400 450 500

Fr'equenc q {Hz}

Configuration 4

ORSPL - 94.3 dB

ALEVEL = 80.7 dB(R}

Configuration 5

ORSPL - 95.4 d8

RLEVEL = 8I.S dBlR)

Con|ig.ration6

0RSPL - 96.5 d8

RLEVEL - 81.2 dBlR]

Configuratiol_ 7

ORSPL - 98.0 dB

RLEVEL - 82.2 dBIAI

FIGURE 4-14. Measured Interior Noise Spectra in Row 6 for the Seat

with the Maximum A-Level (Configurations 4, 5, 6 and 7).

53

0

/IIIIII

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6 'uo!loJalaaa VgO'O 00"0

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56

5 Analysis of Quiet Cabin Noise Levels

The fully treated Quiet Cabin configuration (configuration 4) was tested during 8x8

Flight 1955. The test was conducted at 35,000 ft and at a Mach number of 0.76. To

investigate the noise characteristics of this configuration, data were acquired at each

of the following rotor speeds: 1170, 1200, 1225, 1250, 1260, 1270 and 1280 rpm. The

BPFs associated with these speeds are 156.0, 160.0, 163.2,166.4, 168.0, 169.6 and 171.2

Hz, respectively.

5.1

5.1.1



Since the BPF is the dominant propeller noise contributor to the interior noise (see

Figure 4-11), the variation of the BPF tone level with propeller speed (at constant

aircraft Mach number) was reviewed. As shown in Figure 5-1, the BPF level measured

in the forward propeller plane of rotation generally increases with increasing propeller

speed, while tlle level in the aft plane of rotation does not vary significantly with

propeller speed.

It is interesting to observe that in the forward plane noise levels are highest at the

lower microphone, and in the aft plane noise levels are highest at the upper microphone.

This is primarily the result of the dynamic effects of the rotational direction of the

blades, combined with the shielding effects of the fuselage curvature. Looking forward

from the left side of the aft fuselage, the forward rotor turns counterclockwise towards

the bottom of the fuselage, while the aft rotor turns clockwise towards the top of the

fuselage. Thus noise levels are highest where the blades advance towards the fuselage,and lowest where the blades recede.


Earlier, it was shown that the three dominant components of the pylon vibration

are the tones at the BPF, 2BPF and UN1 frequencies. Figure 5-2a presents the data

acquired on the aft engine mount in the vertical, lateral and longitudinal directions and

Figure 5-2b the data acquired on the inboard side of the pylon, on each of the two spars

in the vertical direction. The BPF data acquired on the engine mount in tile lateral

direction exhibits a strong dependence on the propeller rotational speed with a peak at

1250 rpm. For all practical purposes, the 2BPF response at the engine mount is very

low. The UN1 tone also exhibits a propeller speed dependence but is characterized by

57

a generally decreasing amplitude with increasing speed. Also, the lateral response for

UN1 is weaker than the longitudinal and vertical responses.

The behavior of the data acquired on the pylon spars on the inboard side of tile

pylon is different than on the engine mount. The vibration levels at BPF and 2BPF

increase with increasing propeller speed, a trend similar to that shown in Figure 5-I for

the BPF SPL in the forward propeller plane. Acceleration levels on the aft pylon spar

are greater than those on the forward spar for both BPF and 2BPF. The vibration at

the UN1 frequency measured at these locations does not appear to be influenced by

propeller speed; acceleration is higher on the forward spar than on the aft spar.

5.2 Cabin Noise Levels

For these test points interior noise levels were measured in seat rows 4 and 6.

Figures 5-3 and 5-4 present the measured levels in these two rows, respectively. Each

figure presents the A-weighted overall levels along with the corresponding broadband

and tonal component contributions. Also, the data acquired at microphone position

3 (the aisle measurement location) are presented separately from the data acquired at

the other positions (seat measurement locations).

At position 3, it can be observed that the noise level does not vary substantially

with propeller speed. Further inspection of the data reveals that the contribution

of the tones to the overall level is indeed fairly constant. The data acquired at the

row 4 measurement location exhibits a small increase with increasing propeller speed,

however the variation is too small to be clearly distinguished from the expected exper-

imentai scatter. At the seated positions, the maximum measured levels are shown to

increase with increasing propeller speed, and peak near 1250 rpm. Inspection of thetone contribution shows the same trend.

The overall levels measured at each seat row are very similar. However the tone

contribution to the level of seat row 4 is generally 1 to 3 dB higher than that observed

at seat row 6 while the reverse is true for the broadband contribution.

The variation with propeller speed of the maximum amplitude measured for each

major contributing tone is presented in Figure 5-5. It is readily observed that BPF is

the only tone that exhibits a level increase with increasing propeller speed. Another

feature of the data presented is that the JN1 and BPF tone levels are higher in row 4than in row 6.

Figures 5-6 through 5-9 present the variation of the amplitude of each of the major

tones at each seat location of each row. In Figure 5-6, one can observe that the

distribution of the BPF tone amplitude in each seat row does not follow a particular

58

pattern. For example, at 1225 rpm the amplitude of the BPF tone is nearly constant

across row 6, while at 1170 rpm it decreases toward the right side (JT8D side) of the

aircraft and at 1280 rpm it increases toward the same side. It can however be noted

that, on the right side of the aircraft, higher propeller speeds generally result in higher

BPF tone amplitudes. The data scatter is also observed to increase from slightly more

than 5 dB on the UDF side to slightly more than 10 dB on the JT8D side. In row 4,

levels vary more abruptly from seat to seat for each rotor speed. The trend observed

in Figure 5-5 of higher tone levels in row 4 at 1250 and 1260 rpm can be seen in these

data.

In Figure 5-7, the 2BPF data reveal no apparent trends. The levels typically vary

4 to 7 dB across each seat row, except for the small data scatter at locations 1 and 3

of row 4.

Figures 5-8 and 5-9 present the data associated with the UN1 and the JN1 tones,

respectively. Both sets of data exhibit a very small scatter. This might have been

expected as each of the two engine spools were rotating at nearly constant speed, while

the rotor speed was varied among the seven test points. The other striking feature for

both tones is the low levels observed at seat location 2 of row 6, nearly 10 dB lower

than measured at any other seat location. For JN1, seat location 2 of both row 4 and

row 6 exhibit low levels.

Finally, Figure 5-10 presents the amplitude variation of the tone summation in the

same manner. The scatter is generally small (5 dB or less). No particular trend can

be observed, except on the right side of row 6 where higher propeller speed results in

higher tone summation levels, a trend similar to that observed for the BPF tone. Also,

the decrease in level at seat location 2 of row 6 observed for the UN1 and JN1 tones

occurs for the tone summation levels as well.

Comparison of the levels in Figures 5-6 through 5-10 shows that for most seat

positions and propeller speeds, the tone component of the interior noise is dominated

by the BPF tone, followed in importance by the JN1 tone and then the UN1 tone. The

2BPF tone is not a major contributor to the total tonal energy.

The broadband component of the interior noise is shown in Figure 5-11. The broad-

band level is naturally independent of the propeller speed and remains constant across

the cabin. The slightly higher levels at location 3 are expected since they were measured

at a standing location, closer to the air conditioning ducts at the top of the cabin.

The overall level variation shown in Figure 5-12 indicates very little dependence

oi1 the propeller speed as expected from the data presented in Figures 5-3 and 5-4.

Comparison of the levels in Figures 5-10, 11, and 12 indicates that generally the level

of the broadband component is equal to or higher than the level of the tone component,

in agreement with the data presented in Section 4.5.

59

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LEGEND 7

oot o,.o,0oo,o oo.t

.60 1180 1200 1220 1240 _260 1200 1300Forward and Aft Rotor Speed, rpm

Pltnl! '

FIGURE 5-1. Variation of BPF Sound Pressure Level with

Propeller Speed.

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FIGURE 5-3. A-weighted Interior Noise Levels At Row 4 (Configuration 4).

63

MAXIMUM LEVELS AT POSITION 3 ONLY

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1200 1250

ROTOR SPEED, rpm

FIGURE 5-4. A-weighted Interior Noise Levels At Row 6 (Configuration 4).

64

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CONDITIONS

1170 RPMi POINT 7

1200 RPM; POINT 6

1225 RPM; POINT 5

1250 RPM i POINT 4-

1260 RPM; POINT 3

1270 RPM; POINT 2

1280 RPM; POINT I

I

UDF

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SEAT POSITION

6

JT8D

SIDE

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CONDITIONS

o 1170 RPM; POINT 7 BPFo t200 RPM; POINT 6

A 1225 RPM; POINT 5

+ 1250 RPM; POINT 4

x 1260 RPM; POINT 3o 1270 RPM; POINT 2 ,,/__

_280 R_

u)Ic) 1 j I I ..........

I 2 3 4 5 6

SEA1- POSITION

FIGURE 5-6. Variation of the BPF A-weighted Level in Row 4 and Row 6.

66

i..,q

Row 4

GO

OCO

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U9

UDF

SIDE

tf)

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Aisle

CONDITIONS

[] 1170 RPM; POINT 7

o 1200 RPM; POINT 6

a 1225 RPM; POINT 5

+ 1250 RPM; POIHT 4-

x 1260 RPM; POINT 3

¢' 1270 RPM; POINT 2

v 1280 RPM; POINT 1

3 4-

SEAT POSITION

6

JTSD

SIDE

Row 6

Co

oco

<_r--tn

>

CONDITIONS


o 1200 RPM; POINT 6

/, 1225 RPM; POINT 5

+ 1250 RPM; POINT 4-

x 1260 RPM; POINT 3

o 1270 RPM; POINT 2

v 1280 RPM; POINT 1

2BPF

uJ_J1

o

t_t_

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1 2 [3 4 5 6

SEAT POSITION

FIGURE 5-7. Variation of the 2BPF A-weighted Level in Row 4 and Row 6.

67

I..F_

Row 4

(30

OGO

If)

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0D

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CONDITIONS

a 1170 RPM; POINT 7

o 1200 RPM; POINT 6

a 1225 RPM; POINT 5


x 1260 RPM; POINT 3

o 1270 RPM; POINT 2

v 1280 RPM; POINT 1

UN1

i

m

!

6

JTSD

SIDE

Row 6

CONDITIONS

1170 RPM; POINT 7

1200 RPM; POINT 6

1225 RPM; POINT 5

1250 RPM; POINT 4

1260 RPM; POINT 3

1270 RPM; POINT 2

1280 RPM; POINT 1

UN1

I 2 3 4 5 6

SEAT POSITION

FIGURE 5-8. Variation of the UN1 A-weighted Level in Row 4 and Row 6.

68

co

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3

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SIDE

u5CO

OCO

u5

q:r_

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JN1

CONDI]qONS

__ a 1170RPM; POINT7"_, _ _ o 1200RPM; POINT 6

/f A i225RPM; PO,NT 5

"%%.. /11 + 1250 RPM; POINT 4-_ RPM; POINT 3x 1260

o 1270 RPM; POINT 2

Aisle _ 1280 RPM; POINT 1.L .I. L

1 2 3 4 5

SEAT POSITION6

JTSD

SIDE

Row 6

CONDITIONS JN1

a 1170RPM; POINT 7o ]200 RPM; POINT 6A 1225 RPM; POINT 5

+ 1250 RPM; POINT 4-x 1260 RPM; POINT 3o 1270 RPM; POINT 2 -_

J J I

I 2 3 4 5 6

SEAT POSITION

FIGURE 5-9. Variation of the JN1 A-weighted Level in Row 4 and Row 6.

69

uqCO

o

LL_

.3w_

I

tO

0tD

u_u3

Row 4

Tone Summation

Aislet t

[]

O

A

+

X

O

V

I

CONDITIONS

1170 RPM; POINT 7

1200 RPM; POINT 6

1225 RPM; POINT 5

1250 RPM; POINT 4-

1260 RPM; POINT 3

1270 RPM; POINT 2

1280 RPM; POINT 1

1

UDF

SIDE

t_

2 3 4- 5

SEAT POSITION6

JT8D

SIDE

Row 6

CO

OCO

,_ t--

z3

JLJ_

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K)

C)K)

fibt/b

Tone Summation

CONDITIONS


o 1200 RPM; POINT 6

t, 1225 RPM; POINT 5

+ 1250 RPM; POINT 4

x 1260 RPM; POINT 3

o 1270 RPM; POINT 2

v 1280 RPM; POIN 1 1

I I I I

1 2 3 4 5 6

SEAT POSI I ION

FIGURE 5-I0. Variation of tlleTone Summations in Row 4 and Row 6.

7O

Row 4

t_

Broadband

<_ r--

Ill/I

,.O

O_D

Aisle

±LO

1 2 3 4-

UDF SEAT POSITION

SIDE

C)

CONDITIONS

o 1170 RPM; POINT 7

o 1200 RPM; POINT 6a 1225 RPM; POINT 5

+ 1250 RPM; POINT 4.

x 1260 RPM; POINT 3

o 1270 RPM; POINT 2

v 1280 RPM; POINT 1

.L I

m

6

JT8D

SIDE

Row 6

(D

Broadband

o

if3

<r--

Ld__1I

_D

0

u3

CONDITIONS


o 1200 RPM; POINT 6

a 1225 RPM; POINT 5


x 1260 RPM; POINT 3

o 1270 RPM; POINT 2

v 1280 RPM; POINT 1

I 1 1 I

1 2 3 4- 5 6

SEAT POSITION

FIGURE 5-11. Variation of the Broadband Noise Levels in Row 4 and Row 6.

71

(3O

Row 4

o Overall

tt_I

"1

-1:9

w_>LU__Ji

Oi£)

Aislei [

,n CONDITIONS1170 RPM; POINT 7

;o 1200 RPM; POINT 6

,', 1225 RPM; POINT 5


x 1260 RPM; POINT 3

o 1270 RPM; POINT 2

v 1280 RPM; POINT 1

I 2 3 4 5 6

UDF SEAT POSITION JT8D

SIDE SIDE

| I

Row 6

to

t_

ED

J

>t__Ji

O_O

tf_tO

CONDITIONS

a 1170 RPM; POINT 7

o 1200 RPM; POINT 6

1225 RPM; POINT 5


x 1260 RPM; POINT 3

o 1270 RPM; POINT 2

v 1280 RPM; POINT 1

I t I I

3 4

SEAT POSITION

FIGURE 5-12. Variation of the Overall A-weighted Levelsin Row 4 and Row 6.

72

6 Comparison of 10x8 and 8x8 Data

111 this section, comparisons between 10x8 and 8x8 measurements are presented.

The 10x8 data were coUected during Flight 1964, flown at 35,000 ft at a Math number of

0.76. As for 8x8 Flight 1955, the aircraft interior was fully furnished and configured with

the complete Quiet Cabin package of noise control treatments (that is, Configuration

4). The 10x8 data were acquired at nominal rotor speeds of 1240, 1256, 1264 and 1278

rpm. The blade passage frequencies of the forward/aft rotors associated with these

speeds are 207/165,209/168, 211/169, and 213/170 Hz, respectively.

It should be noted that this rotor speed range for the 10x8 flights is much nar-

rower than the range for the 8x8 flights, thereby fimiting comparisons. In addition,

selected exterior and interior transducers were relocated: some fuselage microphones

were moved to the pylon, some pylon accelerometers were moved to fuselage frames,

and the number of interior microphones was reduced and several were relocated either

inside the cabin or elsewhere. While these changes reveal new information, they also

limit direct comparisons between transducers.

6.1

6.1.1



Figure 6-1 compares narrowband spectra measured at the same aft fuselage location

(above the pylon, in the forward rotor plane). In the 8x8 spectrum, tones at BPF(8) and

its harmonics are evident, as are the individual subharmonic tones at approximately 21

Hz intervals. In the 10x8 spectrum there are two sets of BPF tones, for the BPF(10) and

BPF(8) harmonics. In addition, occasional interaction tones appear in the spectrum

(at BPF(10) + BPF(8), for example).

Comparisons of BPF levels versus rotor speed on the fuselage in the rotor planes

are shown in Figures 6-2a and 6-2b. Figure 6-2a shows 8x8 engine BPF(8) levels above

the pylon in the forward and aft rotor planes (from Figure 5-1). Superimposed on the

figure are the 10x8 engine levels measured at the same locations. For this comparison,

the BPF(10) and BPF(8) levels of the 10x8 engine have been added together, since

the 8x8 engine BPF(8) levels are due to the combined effect of both 8-bladed rotors.

However, at both measurement locations the BPF(8) level is significantly higher than

the BPF(10) level for the 10x8 engine, so that the summed level is dominated by the

contribution from the aft rotor. Figure 6-2a shows that the 10x8 levels are 3 to 5 dB

higher in the aft plane than the 8x8 levels, and 1 dB lower in the forward plane.

In contrast, Figure 6-2b shows corresponding data measured on the fuselage below

73

the pylon in the forward plane (note that the two measurement points are shifted

slightly). Here, the 10x8 engine levels, which are dominated by the BPF(10) levels, are

3 to 4 dB lower than the 8x8 engine levels.

It should be noted that at the upper microphone locations the 10x8 BPF levels

are dominated by the noise of the aft (8-bladed) rotor, while at the lower microphone

locations the 10x8 BPF levels are dominated by the noise of the forward (10-bladed)

rotor. As discussed in Section 5 for the 8x8 exterior data, this reflects the clockwise

rotation of the aft rotor towards the top of the fuselage and the counterclockwise

rotation of the forward rotor towards the bottom of the fuselage.

Figures 6-2a and 6-2b indicate that the variation of level with rpm is similar for the

two engine configurations, although the speed range may be too narrow for significant

trends to emerge. For the measurement locations studied, the highest aft fuselage levels

were measured during the 10x8 engine flight tests.

When the 8x8 engine flight tests were conducted, the blade pitch was varied from

test point to test point to maintain constant shaft horsepower for all rotor speeds.

However, during the 10x8 engine flight tests shaft horsepower varied with rotor speed.

Estimates of the relative differences between 10x8 and 8x8 engine BPF levels were made,

taking these shaft horsepower differences into account, as well as the differences in rotor

diameter, blade count, and tip clearance between the two propulsor configurations.

These estimates agree in trend with the measurement results discussed above: higher

BPF levels in the aft rotor plane and lower BPF levels in the forward plane for the 10x8

engine compared to the 8x8 engine, and higher maximum levels on the aft fuselage for

the 10x8 engine.

6.1.2 Pylon and Fuselage Vibration Levels

Figures fi-3a and 6-3b present comparisons between 10x8 and 8x8 pylon accelera-

tions. In Figure fi-3a, the accelerations measured at the aft engine mount are compared.

In the vertical direction, 10x8 accelerations are considerably lower than 8x8 accelera-

tions. In the longitudinal direction, 10x8 accelerations are comparable to or lower than

8x8 accelerations.

In Figure 6-3b, comparisons of 10x8 and 8x8 accelerations measured inboard at the

forward spar show that BPF(8) vertical accelerations are considerably lower for the

10x8 engine, while BPF(10) vertical accelerations for the 10x8 engine are somewhat

higher than 8x8 BPF accelerations in the lower portion of the rpm range, and somewhat

lower than 8x8 BPF accelerations in the higher portion of the rpm range. The vertical

accelerations at UN1 are comparable for the two systems.

Summarizing, these figures show that 10x8 vibration levels in the pylon are typically

74

the same or lower than 8x8 levels.

Finally, Figure 6-4 compares 10x8 and 8x8 acceleration levels measured on the

fuselage in the forward rotor plane; the 10x8 levels are a summation of the BPF(10)

and BPF(8) levels, but are dominated by the BPF(8) levels. The variation with rpm is

similar for the two systems. However, the figure shows that 10x8 levels are higher than

8x8 levels; this observed trend is opposite to the trend in noise levels at the nearby

fuselage location (shown in Figure 6-2a).


Figures 6-5a through 6-5e present comparisons between 10x8 and 8x8 cabin noise

levels, for each of five seat positions in the aft cabin where levels were monitored in both

sets of tests. On each figure, the overall A-level and broadband and tonal components

measured during 10x8 and 8x8 flights over the same range of rotor speeds are presented.

The figures show that overall levels for the two systems are comparable at positions

1 and 3 in row 6 and position 1 in row 4, and are lower (by about 2 dB) for the 10x8

configuration at position 4 in row 6 and position 3 in row 4. Broadband levels are

comparable for the two engine configurations at each seat; the differences in overall

levels at the latter two positions are due to lower tone levels during the 10x8 flights.

At position 4 in row 6, the tone component is 2 to 3 dB lower for the 10x8 configura-

tion, while at position 3 in row 4 the 10x8 tonal component is 5 dB lower. Figures 6-6a

and 6-6b compare the 10x8 and 8x8 spectra measured at these positions for comparable

rotor speed. The JN1, JN2, and UN1 tones are seen to be roughly comparable in level

between the 10x8 and 8x8 spectra at both positions, but the 10x8 BPF tones are lower

(and at position 3 in row 4 are much lower) than the 8x8 BPF tones.

For the 10x8 engine, exterior acoustic levels at the blade passage frequencies were

found to be higher but pylon vibration levels were found to be comparable to or lower

than the corresponding levels of the 8x8 engine for the same frequencies. The reduction

in interior BPF noise levels in spite of the increase in the exterior BPF acoustic loads

suggests that these cabin tones are due to the structureborne propagation of pylon

vibrations through the fuselage.

75

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88 (.-. __

7 Comparison of Cruise and Non-Cruise Data

In this section, comparisons between cruise and non-cruise measurements are pre-

sented, for both 8x8 and 10x8 engine configurations. "Cruise" refers to the standard

high altitude, high speed cruise conditions of 35,000 ft. and Mach 0.76 at which all

of the data reported previously were obtained. "Non-cruise" refers to flight conditions

at lower altitudes and at the same or lower Mach numbers, although the aircraft is

operating in a cruise mode at these non-standard conditions.

Only a limited number of test points containing interior noise data are available

for the non-cruise conditions. Table 7-1 lists the non-cruise test points included in

this analysis, and the corresponding cruise test points used for comparison. These

test points were selected to have overlapping rotor speed ranges to the extent possible,

and to have the Quiet Cabin (or subsequent) treatment configuration. To facilitate

comparisons, cruise conditions are labelled with a "C", and non-cruise conditions are

labelled with an "N".

7.1 Acoustic and Vibratory Loads

7.1.1 Exterior Noise Levels

Figure 7-1 compares 8x8 BPF levels for four flight conditions measured at two

locations on tile fuselage exterior in the forward propeller plane. The figure shows a 2

to 4 dB increase between cruise levels (Condition C) and the levels at Condition N3,

but nominally no difference between cruise levels and the levels at Conditions N1 and

N2.

Estimates of the expected differences in BPF tone levels between the cruise and

non-cruise conditions were made, based on empirical adjustments for differences in

altitude (due to pressure and speed of sound differences) and helical tip Mach number.

For Condition N3, the estimated difference in level is 3.5 dB; for Conditions N1 and N2

the estimated difference is -0.5 dB. These agree well with the measured exterior data

shown in Figure 7-1.

In Figure 7-2, 10x8 BPF(10) and BPF(8) levels for three flight conditions are com-

pared, at the upper measurement location. The levels for Condition N5 are 7 to 11 dB

lower than the cruise levels for both BPFs. The single data point for Condition N4 is

higher than cruise levels by 2 dB for BPF(8), and lower than cruise levels by about 4

dB for BPF(10).

The estimated difference between the BPF levels for cruise conditions and Condition

89

N5 is -7.5 dB, again in agreement with the measured data. However, for Condition N4

the estimated difference in levels relative to cruise levels is -0.5 dB. The BPF levels for

the single data point in Figure 7-2 for this condition is at variance with this estimate.


Vibration levels on the pylon are available for the 10x8 flights only, and only at

the rear engine mount in the longitudinal and vertical directions. Figures 7-3 and

7-4 compare vibration levels measured in these two directions, respectively, for the

BPF(10), BPF(8), and UN1 tones for the three 10x8 flight conditions.

In the longitudinal direction, Figure 7-3 shows that the levels of the three tones for

the cruise conditions and Condition N5 are generally comparable, depending on rotor

speed. The single point for Condition N4 varies inconsistently in comparison with the

other data, as was the case for the exterior noise levels for this data point.

In Figure 7-4 it can be seen that in the vertical direction, BPF(8) vibration levels

for cruise conditions are higher than for Condition NS, while the opposite is true for

BPF(10) levels. UN1 levels tend to overlap for these two conditions. Again, the tone

levels for the single data point for Condition N4 are at variance with the levels for theother conditions.


The maximum BPF tone levels measured among several aft cabin seats are shown

in Figure 7-5, for the various flight conditions. In the top of the figure, the maximum

8x8 BPF levels measured among seats 1, 3, 4, 5, and 6 in row 6 are compared. The

general trends of the data in this figure reflect the exterior levels shown in Figure 7-1:

higher tone levels compared to cruise by 2 to 4 dB for Condition N3, and comparabletone levels for Conditions N1 and N2.

Changes in the BPF tone levels measured in the aft cabin for non-cruise conditions

relative to cruise conditions for the 8x8 flights thus appear to be directly related to cor-

responding changes in the exterior acoustic loads. This relationship is not as clearcut,

however, for the 10x8 flights. BPF(10) and BPF(8) levels are shown in the middle and

bottom portion of the figure, respectively, and represent the maximum levels measured

among seats 1, 3, and 4 of row 6 and seats 1 and 6 of row 5. The decrease in level

of 7 to 11 dB from cruise conditions to Condition N5 observed in the exterior data of

Figure 7-2 for both BPF tones is not clearly reflected in the interior noise levels. For

BPF(10), this trend does occur for the lower rotor speeds, but the difference decreases

significantly for the higher rotor speeds. For BPF(8), the levels for the two conditions

9O

tend to overlap. These relative levels also appear unrelated to the measured pylon

vibration levels shown in Figures 7-3 and 7-4. Finally, the BPF levels for Condition N4

also vary differently compared with cruise levels than either the exterior noise data or

the pylon vibration data.

It is reasonable to expect that the maximum BPF tone levels measured in the cabin

vary in accordance with the maximum BPF tone levels measured on the aft fuselage, as

was seen to be the case for the 8x8 data. For the 10x8 data, the only available exterior

microphones were not those at which the maximum near field levels were expected

to occur, based on the data presented in Section 6. Thus the discrepancies observed

between the exterior and interior data for the 10x8 flights may be a result of using

inappropriate exterior levels.

Figures 7-6 and 7-7 present the interior noise levels for the 8x8 and 10x8 flights,

respectively, in terms of the overall A-levels, and the tone and broadband components

of the A-level. Each figure consists of several pages, one for each seat position.

The overall A-levels for the 8x8 flights depicted in Figure 7-6 show typically a 5 dB

increase for all seats for Condition N3 as compared with levels under cruise conditions.

For Conditions N1 and N2, overall levels are 0 to 5 dB higher than for cruise conditions,

averaging about 3 dB higher. In each case, these increases are the result of higher

broadband levels, which dominate the interior levels for the non-cruise conditions. The

estimated change in boundary layer noise levels for the non-cruise conditions relative

to cruise conditions is a 5 dB increase for 22,500 ft altitude and Mach 0.77 (Condition

N3), and a 3 to 4 dB increase for 22,500 ft altitude and Mach 0.69 to 0.70 (Conditions

N1, N2). These expected changes in exterior levels are reflected in the observed changes

in broadband interior levels measured at each of the aft cabin seats. Although BPF

tone levels also increase for non-cruise conditions (as seen in Figure 7-5), the larger

increases in broadband noise levels have the more significant effect on the overall cabinnoise levels.

For the 10x8 flights, Figure 7-7 shows that in seat row 6 overall A-levels decrease

by 1 to 3 dB for Condition N5, relative to levels for cruise conditions. As for the 8x8

data, this change reflects the corresponding change in broadband levels in seat row 6.

The estimated change in boundary layer noise level is a 1.5 dB decrease for these non-

cruise conditions relative to cruise conditions; the measured interior broadband levels

are in good agreement with this prediction. For seat row 6 measurements, the levels of

the broadband component exceeds the levels of the tone component, especially for the

non-cruise data where the tone component has been significantly reduced, resulting inlower overall levels which track the lower broadband levels.

The seat row 5 data for the 10x8 configuration also show lower broadband levels

for these non-cruise conditions, but the tone levels are higher in this row than in row

91

6. The net result is that overall levels in seat row 5 for these non-cruise conditions

are comparable to (and for some rotor speeds higher than) the overall levels for cruise

conditions.

As noted for the 10x8 BPF data, the levels measured during Condition N4 varied

considerably, and were inconsistent among locations. Similarly the interior noise levels

for this condition are also variable, and do not follow expected trends.

92

TABLE 7-1. Test Point Summary for Non-Cruise Comparisons

RUN l UHBNUMBER CONFIG.

19550A01

19550A02

19550A03

19550A04

19550A05

19550A06

19550A07

19640A01

19640A02

19640A03

19640A04

19640A05

19640A06

19640A15

19640B01

19640B02

19640B03

19640B04

20330K01

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8x8

ROTOR SPEED

(RPM)

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1225

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1200

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1265

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ALTITUDE

(ft.)

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22500

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0.76

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C

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108

8 Identification of Major Transmission Paths

When exterior acoustic and mechanical forces excite a complex structure, determi-

nation of the major transmission paths by direct measurements is generally not feasible.

Several analysis methods are available, however, which provide a partial definition of

transmission paths. For example, various techniques to compare measured interior and

exterior noise and vibration levels can be used to infer the relative strengths of specific

paths. Measured sound intensity data can be used to identify major radiating surfaces

and thereby aid in the interpretation of other measurement results. Partial coherence

techniques can be used to examine the degree to which individual signals measured in

the interior or along specific transmission paths are related to selected source signals,

to further deduce transmission path strengths. Taken together, the results of these var-

ious analyses can provide a complete understanding of the transmission characteristics

into the aircraft cabin.

In this section, all of these analysis techniques are utilized to determine the major

transmission paths on the Demonstrator. Prior to the flight test program, the three

paths shown in Figure 4-1 were predicted to be the major paths. The various analyses

of measured flight test data have focused on developing a better understanding of the

transmission characteristics on these, and other, paths.

In the first subsection below, the results of a sound intensity survey conducted

on a dedicated 8x8 flight are presented. In the following three subsections, each of

the transmission paths shown in Figure 4-1 is examined using a variety of interior

and exterior data measured during 8x8 flight 1955. The relative strength of each

path is assessed using these data and the sound intensity survey results. Finally, the

last subsection presents the results of partial coherence analyses of selected data from

8x8 flight 1955 and 10x8 flight 1964, which provide an independent assessment of

transmission path strengths.

8.1 Results of Sound Intensity Survey

Figures 8-1 and 8-2 show sound intensities and sound power levels, respectively,

for various scanned surfaces in the passenger cabin measured during 8x8 flight 2026.

Measurement locations are defined in Figures 3-7 and 3-8. It should be noted that for

these sound intensity measurements, the fuselage configuration (configuration 5) was

different from the Quiet Cabin configuration (configuration 4). In addition to removal

of the engine dynamic absorbers, several items were removed from the cabin to facilitate

the survey: the seats, the carpeting on the floor, the carpeting on the left engine pylon

bulkhead, and the section of the double wall bulkhead over the aft pressure bulkhead.

109

On each figure, the measured levels are listed for the four one-third octave bands

of 100, 160, 200, and 315 Hz; these bands contain the 3N1, BPF, UN1, and 2BPF

tones, respectively. Positive values indicate energy flow from the scanned surface, while

negative values indicate energy flow into the surface. Levels in parentheses indicate

that the pressure intensity index is greater than 20 dB, and therefore the data are

questionable.

For all four tones of interest, the highest intensities and power levels arise from the

aft bulkheads (both engine pylon bulkheads and the aft pressure bulkhead) and the

right upper and lower sidewalls (in the aft part of the cabin). There are also significant

but lower levels from the left sidewall (for UN1) and from the aft floor (for BPF).

8.2 Cabin Sidewall Path

Figure 8-3a shows exterior spectra for several locations on the fuselage for 8x8

flight 1955, test point 0A03 (with a rotor speed of 1260 rpm). While tones at the BPF,

2BPF, and subharmonics are very prominent in the vicinity of the propeller planes

of rotation, these tones rapidly diminish as the location moves forward. Outside the

passenger cabin the tones have nearly completely disappeared, with levels at or below

the broadband noise (due to the boundary layer).

Interior spectra measured in the last seat row, as shown in Figure 8-3b, depict strong

tones at the BPF and 2BPF superimposed on a broadband background (also due to

the boundary layer). If these BPF and 2BPF tones measured in the cabin were the

result of transmission through the cabin sidewall, their levels would be at or below the

broadband levels, corresponding to the spectral characteristics of the acoustic loads on

the fuselage exterior. Since this is not the case, it is clear that the tones in the interior

are not the result of direct transmission through the cabin sidewall.

To further support this conclusion, the power levels presented in Figure 8-2 show

that the left sidewall is not a major source for tones at the BPF and 2BPF frequencies

(note the higher power levels from the right sidewall at these frequencies). Thus the

tones observed in the cabin are not transmitted along the airborne path from the UHB

engine through the cabin sidewall.

The broadband background noise in the cabin, however, is hkely the result of trans-

mission of exterior boundary layer noise through the cabin sidewall, as is the case

on most conventional aircraft. In Section 7, estimates of the differences in exterior

boundary layer noise levels for several flight conditions were in good agreement with

measured differences in interior broadband levels for the same conditions. This agree-

ment tends to verify the concept that interior broadband levels are tile result of the

exterior boundary layer noise.

110

8.3 Aft Section Path

Figure 8-4 shows spectra measured at two locations just behind the aft pressure

bulkhead. This figure shows major tones at the BPF and 2BPF, with maximum levels

on the order of 110 dB and 100 dB, respectively (note the absence of engine rotational

tones in these spectra). For comparison, the interior noise spectra of Figure 8-3b show

naaximum BPF levels on the order of 90 dB, and maximum 2BPF levels on the order

of 70 dB. If the tones in the cabin are the result of transmission through the pressure

bulkhead, the bulkhead is providing a reduction in level of about 20 dB at the BPF,

and about 30 dB at 2BPF.

Independent measurements were conducted in a Douglas Aircraft Company ground

test facility using a DC-9 test article and a loudspeaker located in the aft section as a

source (see NASA CR 181819). With the same double wall treatment apphed to the

pressure bulkhead, the reduction in level between a corresponding location behind the

bulkhead and a typical seated position in the aft seat row was on the order of 35 dB

for the frequency range corresponding to the BPF, and on the order of 45 dB for the

frequency range corresponding to 2BPF. Thus, if the cabin tones were strictly a result

of transmission through the pressure bulkhead, their levels would be approximately 15

dB lower than the levels actually measured.

The ground measurements also showed that with the double bulkhead removed, the

pressure bulkhead still provides about 35 dB noise reduction for the frequency range

corresponding to the BPF. This would tend to confirm the results reported in Section

4 for Treatment 7, i.e. that interior BPF levels were relatively unchanged with the

removal of the section of the double wall bulkhead over the pressure bulkhead door.

For higher frequencies (typically above 200 Hz), the ground measurements showed

an improvement in noise reduction due to the double wall treatment of 10 to 20 dB,

generally increasing with frequency. This improvement was also observed in the UHB

Demonstrator measurements. For example, Figure 8-5 compares spectra at the same

seat location for the flight tests corresponding to the Quiet Cabin configuration (treat-

ment configuration 4), and treatment configurations 6 and 7. In this figure, the spectra

are plotted to 2000 Hz. The presence of the tones at the higher multiples of the BPF

when the double wall bulkhead was removed (configuration 7) is clear, indicating that

this path would be important for these higher frequencies if the bulkhead were not

treated. Although these tones do not contribute significantly to the A-level in the

cabin, they would be perceptible to passengers. Note also the increasing level of the

2BPF tone from configurations 4 to 7, as observed earher in Figure 4-11.

The discussion above has shown that the transmission path from the aft unpressur-

ized section through the pressure bulkhead into tile cabin is not significant, particularly

when the pressure bulkhead is treated with a double wall. Nevertheless, it is interesting

111

to examine the first portion of this "aft section path", i.e. the transmission of sound

(and particularly BPF tones) from the UHB engine into the aft unpressurized section.

At first thought, it would appear that acoustic forces from the two blade rows would

impinge on the aft fuselage, and then transmit directly through the fuselage wall into

the aft unpressurized section. Analysis of measured data, however, shows that this is

not exactly the situation.

Figure 8-6 compares BPF levels measured at several transducers as a function of

UDF rotor speed for flight 1955. The data are plotted for selected fuselage exterior

microphones in the two propeller planes, for a fuselage accelerometer in the forward

prop plane, and for two microphones in the aft unpressurized section. The similarity

between the levels at both aft section interior microphones and the fuselage accelerom-

eter levels is easily seen. In fact, the computed linear regression between the levels at

the centerline interior microphone and the accelerometer levels has a correlation coem-

cient of 0.98. Thus, it appears that the BPF levels in the aft unpressurized section are

directly related to (and caused by) the structural motion of the fuselage.

In comparing the accelerometer levels to the exterior microphone levels, there are

dissimilar variations with rpm. A linear regression of the accelerometer levels with

any of the microphone levels shows small correlation. However, the primary area of

disagreement is at 1250 rpm, where accelerometer levels do not rise with increased

acoustic levels (as measured by the two microphones in the forward prop plane). If this

test point is (temporarily) neglected, the agreement improves. The computed linear

regressions between the accelerometer levels and the microphone levels in the forward

prop plane for the remaining six test points have correlation coefllcients of 0.84 for the

upper microphone, and 0.96 for the lower microphone.

The exterior microphones measure the acoustic loads impinging on the fuselage

(generated by the dynamic loads on the rotating propellers, which may include steady

and unsteady components), while the accelerometer measures the structural response

of the fuselage to these loads. This response is not necessarily linear with frequency,

particularly given the complex vertical and canted frame structure and the use of

damping treatment in this area. Conceivably, there is a dip in the structural response

at 166 Hz, the BPF corresponding to 1250 rpm. This nonlinear behavior could explain

the dissimilarity in the accelerometer/microphone data at the 1250 rpm test point.

In summary, the BPF levels in the aft unpressurized section are likely the result of

the structural response of the fuselage to the impinging acoustic loads. With the aft

pressure bulkhead treated (with a double wall), these levels do not propagate signifi-

cantly into the passenger cabin.

112

8.4 Structural Path

As has been discussed earher, the noise levels measured in the cabin consist of a

broadband component due to boundary layer noise, and a tone component which is

comprised of the BPF, 2BPF, JN1, and UN1 tones superimposed on the broadband

spectrum. In the previous sections it was observed that the broadband component

arises from external boundary layer noise transmitted through the cabin sidewall, and

that the various tones did not result from transmission through either the cabin sidewall

or the aft pressure bulkhead. In the following, it will be demonstrated that these tones

are transmitted structurally through the engine pylon and fusdage.

The sources of the tones at the engine rotational frequencies are mechanical forces

within the engines which do not generate airborne noise but which propagate struc-

turally. This is substantiated by review of the noise spectra for the exterior fuselage

microphones (Figure 8-3a), in which the UN1 tone is absent, and review of pylon vi-

bration data (Figure 5-2) which shows a strong UN1 component at both the engine

mount and pylon spar locations.

Indeed, the sound intensity measurements (Figures 8-1 and 8-2) show that for both

UN1 and JN1 tones, the aft bulkheads are major radiating surfaces, as are the rear

sidewalls. It appears that the JN1 tone propagates through the JT8D pylon into

the fuselage, through the pressure bulkhead, the right pylon bulkhead and the right

sidewall. The UN1 tone similarly propagates through the UDF pylon into the fuselage,

through the pressure bulkhead, the left pylon bulkhead and the left sidewall. However,

the UN1 tone also propagates through the entire pylon bulkhead structure to the right

side of the aircraft, and into the right sidewall.

It should be noted that in order to support the UDF pylon, a new left pylon bulk-

head was installed which is significantly stiffer and more massive than the right pylon

bulkhead. Structural energy from the UDF pylon thus has a ready transmission path

through this structure to the other side of the aircraft. It should also be remembered

that the double wall bulkhead section over the aft pressure bulkhead door was removed

during the sound intensity measurements. Although high sound power levels are shown

in Figure 8-2 for the aft pressure bulkhead, it is likely that the sound power flow from

this bulkhead is reduced in the fully treated cabin configuration.

For the BPF and 2BPF tones, the sound intensity measurements again show that the

major radiating surfaces are the aft pressure bulkhead, the two engine pylon bulkheads,

and (for the BPF tone) the right sidewall in the rear of the cabin. Since it was shown

above that these tones do not propagate along an airborne transmission path through

the pressure bulkhead, the sound power flow from the bulkhead must be structurally

induced. Also, the presence of sound radiation at the BPF and 2BPF frequencies from

the right side of the aircraft can only be structurally generated. As a final point, Figure

113

5-2b shows that the vertical acceleration levels measured at the inboard pylon spar for

the BPF tone follow the same trend with rpm as the exterior microphone levels (Figure

8-6) in the forward prop plane. The computed linear regression of the acceleration levels

with tile lower microphone levels has a correlatiou coefficient of 0.97. Thus the BPF

tones propagating through the fuselage are likely caused by the impinging acoustic

loads on the UDF pylon.

In Section 5.2, the cabin noise level distribution for the various tones was discussed.

The sound intensity data provide an explanation for some of the observed trends. For

example, the BPF tone measured in the last seat row appeared to be higher on the right

(JT8D) side than the left (UDF) side for most of the engine rpm settings. Since the

pressure bulkhead had a double wall treatment and the left engine pylon bulkhead was

covered by a coat closet during the flight tests, thereby greatly reducing the acoustic

energy from these two surfaces, the high sound power levels from the right engine pylon

bulkhead and right sidewall would explain the higher measured levels on the right side

of the cabin.

The JN1 tones were also observed to be higher on the right side. Figure 8-2 shows

that the greatest sound power at this frequency emanates from the right sidewall, as

might be expected.

In contrast, the sound intensity results show radiation from both sidewalls and the

aft bulkheads into the cabin for the UN1 tone, and primarily from the aft bulkheads

for the 2BPF tone. These two tones show roughly the same levels on both sides of the

cabin.

In summary, all the tones observed in the cabin appear to be generated by en-

ergy that is propagated structurally, resulting in acoustic radiation from various cabin

surfaces. The major surfaces include all the aft bulkheads and the right rear cabin

sidewall.

8.5 Results of Partial Coherence Analyses

Partial coherence analyses were conducted for 8x8 flight 1955 test point 0A03 and

10x8 flight 1964 test point 0B03. The purpose of these analyses is to identify the major

sources and paths for the BPF tones measured in the cabin. An overview of the analysis

technique and major results are described in this subsection. The mathematical basis

of the partial coherence analysis technique is contained in Appendix B.

114

8.5.1 Partial Coherence Analysis Concepts

The degree to which two signals are related can be measured by the "ordinary

coherence function". A coherence value of 0.1 or less between two signals is low enough

for an assumption of hnear independence, while a value of 0.9 or higher indicates that

the signals are highly dependent. The coherency between two signals, however, is not

an indication of a cause and effect relationship. For example, when there is a single

periodic source impinging on a structure, the measured signals throughout the structure

are all likely to be highly coherent; nevertheless one cannot draw conclusions about

transmission paths from these coherence values alone. Thus the coherence between an

interior microphone signal and a trim panel accelerometer signal may be nearly 1, but

the sound at the microphone is not necessarily caused by the vibration of the trim panel,

nor is the trim panel vibration necessarily caused by sound in the cabin. Similarly if

the periodic source in this example is a rotating propeller generating acoustic energy,

the signals at nearby exterior microphones and accelerometers will also likely be highly

coherent with each other and with the signals at the two interior transducers, but the

presence of airborne and/or structureborne transmission paths cannot be deduced from

these coherence values.

In this example, the signal at the exterior microphone could be considered a measure

of the noise source, while the signal at the interior microphone could be considered a

measure of the response to this source. In other terms, the aircraft structure could

be considered a system with inputs and outputs; the source signal would be an input,

and the response signal would be an output. When there are multiple sources (such as

dual propeller rotors generating multiple acoustic sources and mechanical forces in the

engine generating vibration sources), the noise measured at a specific interior location

can be modeled as a multiple-input single-output process.

A "partial coherence" analysis involves selectively removing the effects of individ-

ual inputs from the output. The partial coherence function is then a measure of the

coherence between a single input and a single output, after removing the effects of spec-

ified inputs. The process of removing these specified inputs from the output is called

"conditioning". The partial coherence analysis begins with the output spectrum. This

spectrum is then "conditioned" as the effect of the first selected input is removed. This

conditioned spectrum is then further conditioned as the effect of the second selected

input is removed, and so on.

If this technique is applied to a specific tone in the output spectrum, then as each

input is removed the level of the tone in the conditioned spectra will sequentially

diminish, for those inputs which have a real effect on the output. That is, if there are

three inputs but only the first two of them affect the output, the conditioned spectra

will show reductions in level only when these two inputs are removed. Conditioning

115

the input with the third output will produce no reduction in level.

There are several subtleties implicit in the use of this technique, some of which

were uncovered during the analyses described below. First, the tone level will decrease

to a background noise floor as the conditioning process occurs, if the set of selected

inputs completely reflects the noise source(s). In practical apphcation the number and

location of transducers may not be sufficient for complete measurement of the source,

resulting in a residual level remaining after the output is conditioned with all the inputs.

Second, for a fixed set of inputs, the overall results of the conditioning process were

found to be independent of the ordering of the inputs. That is, the final conditioned

output spectrum does not depend on the order of removal of the inputs; this is in

contrast with current thinking which holds that the input order is important to the

final conditioned output. However, the intermediate conditioned spectra are highly

dependent on input ordering, and thus the relative importance of each input to the

output cannot be determined from the intermediate spectra alone. Third, if the inputs

are independent of each other the relative reduction in level as each input is separately

removed from the output will provide a rank ordering of the importance of each to the

output. This however is not always the case, since the signal measured at a specific

input transducer may include the effect of multiple sources. Finally, it is critical to

identify inputs and outputs properly. Often transducers selected as inputs are actually

intermediate outputs, resulting in misleading conclusions (particularly in systems with

high coherency among all transducers).

8.5.2 Analysis of 8x8 Data

Figure 8-7 shows the transducers used in the partial coherence analysis of the 8x8

flight test data. The five interior microphones in seat row 6 are used as output trans-

ducers. For input transducers, the four exterior microphones in the two rotor planes (3,

4, 12, 13) are used to represent acoustic sources, while the four accelerometers on the

engine mount (4, 5X, 5Y, 5Z) are used to represent vibration sources. Accelerometers

1, 2, and 3 and aft section microphone 5 are considered to be (intermediate) output

transducers, that is, they measure the response of the fuselage to either acoustic or

vibration excitation, rather than the source of the excitation itself. (In this analysis,

tile cabin sidewall path is ignored since the noise spectra measured on the fuselage

outside the cabin do not contain BPF tones.)

As a starting point, Table 8-1 lists the ordinary coherence between the various

inputs and each cabin output. It is surprising to note that, with a few exceptions,

the coherence values are relatively low. Further, there is considerable variability in

the coherence between a single input and several outputs. For example, the coherence

values between exterior microphone 3 (which directly measures the acoustic energy from

116

the rear propeUer) and interior microphones1, 4, and 5 lie between0.6 and 0.8, whilethe coherenceswith interior microphones3 and 6 aresignificantly lower. Similar trendsare seenin the accelerometercoherencevalues. This may imply a complex interactionamong the various acoustic and vibration sourceswith regard to the interior noiseenvironment.

The results of conditioning eachoutput with the various inputs (both individuallyand sequentially) are shown in Table 8-2. The table hsts the reduction in the level ofthe BPF tone at eachof the five interior microphones,as a result of the conditioningprocess.The first four entries in the table showthe effecton the output microphonesofthe removalof eachinput microphone,individually. The reductions seemto correspondwith the coherence values in Table 8-1; removal of inputs with high coherence result in

large reductions, while removal of inputs with low coherence results in small reductions.

In contrast, the next set of entries in the table show the effect on the output micro-

phones of the sequential removal of the same input microphones, in the order hsted.

Thus for output microphone 1, removal of the effects of input microphone 4 reduces the

BPF level by 1.5 dB (as was also seen in the first entry in the table), but the subsequent

removal of the effects of input microphone 13 further reduces the BPF level by 13.5

dB, for a total reduction of 15 dB. Although the individual effects of conditioning the

output with these two inputs are small (1.5 and 0 dB), the combined effect is substan-

tial. This implies that the two input microphones are measuring different sources (or

different combinations of two sources), and both of these sources are contributing to

the level at output microphone 1.

It is likely that these two sources are in fact the two propellers. Although the

frequency of the tones generated by these propellers is the same (within the resolution

of the spectral analysis), at any point on the fuselage there is a phase difference between

the signals received from each blade row, which varies with location. This varying phase

shift arises from the fact that the propellers are counter-rotating, and at different

distances from the fuselage. Thus there are two separate acoustic sources, of the same

frequency, impinging on the aft fuselage; the level in the cabin is affected by both.

Returning to Table 8-2, it can be seen that for the same output microphone, further

conditioning with input microphones 3 and 12 produces no further reduction in level.

It would appear that two exterior microphones (4 and 13) are sufficient to include the

effects of the two exterior sources.

The same trends observed for output microphone 1 can be observed for the other

output microphones, to varying degrees.

The next portion of the table contains the results of conditioning the output with

the accelerometer inputs. As for the microphone inputs, the accelerometer inputs

exhibit the characteristic that the individual reductions correspond to the coherence

117

values. In contrast to the microphone inputs, however, the sequential reductions for

the accelerometer inputs are more in line with reductions that would be expected based

on the individual reductions.

Finally, the reductions in level at the output microphones can be compared for three

sets of conditioning inputs: all microphone inputs, all accelerometer inputs, and all

microphone and accelerometer inputs combined. The table shows that for each output

microphone, conditioning with all the inputs produces the greatest reductions; these

reductions for all the outputs eliminate the BPF tone from the conditioned spectra

(within 1 dB). This implies that all the acoustic energy at the output is accounted for

by the inputs used in the analysis, and further that both sets of inputs (microphone and

accelerometer) are needed. Thus it can be concluded that the interior BPF levels result

from both acoustic and vibration excitation of the structure. Furthermore, the relative

magnitudes of the reductions produced separately by the microphone and accelerometer

inputs give an indication of the relative strengths of the two types of sources: for output

microphones 1, 3, and 5 the acoustic sources appear to contribute more to the interior

level than the vibration sources, while the reverse occurs for output microphones 4 and

6.

Since the source of the BPF tones in the cabin is a combination of acoustic and

vibration excitation, the issue of transmission path identification still remains. Is the

acoustic energy transmitted solely along the aft section path and the vibration energy

transmitted solely along the structural path, or is there a more complex propagation

scenario? To investigate this further, accelerometers 2 and 3 were analyzed as outputs.

Each was conditioned with the various exterior microphone and accelerometer inputs

used previously. For both outputs, the results of the partial coherence analysis showed

that the BPF levels arose from a combination of acoustic and vibration excitation. Thus

the structural path, through the pylon and fuselage structure, transmits both acoustic

energy (from acoustic loads impinging on the pylon) and vibration energy (from engine

mechanical forces) into the cabin.

In the unpressurized section behind the pressure bulkhead, aft section microphone

5 was also conditioned with the various exterior microphone and accelerometer inputs.

Conditioning this output with these several inputs together only partially reduced the

BPF tone level. Since a relationship between the BPF levels at this microphone and

at exterior accelerometer 1 was observed in Section 8.3, this accelerometer was also

used as an input to further condition the aft section microphone 5 output (despite the

fact that this accelerometer is considered to be an output transducer itself). Again the

level at this output microphone was not fully reduced to the background. The presence

of residual energy at this output after conditioning with these inputs implies that

other sources may affect the BPF levels in the aft section. For example, radiation of

the pressure bulkhead (due to propagation of structural energy through the pylon spars

118

and fuselage structure into this bulkhead) may contribute to the aft section microphone

levels. It should be noted that among the five microphones in this aft unpressurized

section (see Figure 3-3), the BPF level is highest at two microphones: microphone

1 adjacent to the fuselage wall near the forward propeller plane and microphone 5

adjacent to the pressure bulkhead. This supports the hypothesis that vibration of

fuselage structural members contributes to the BPF levels measured in the aft section.

From this discussion it can be inferred that the aft section path is not significant in

the transmission of acoustic energy into the cabin. It would appear, then, that acoustic

and vibration excitation of the pylon is transmitted into the fuselage structure, and

then propagated into the passenger cabin. Structural energy is also propagated into

the aft unpressurized section. These results are in agreement with the conclusions

developed independently in Section 8.3 and 8.4.

8.5.3 Analysis of 10x8 Data

The partial coherence analysis of the 10x8 flight test data was conducted in a similar

manner to that of the 8x8 data. There were, however, some differences in transducers.

For outputs, interior microphones in row six were available for seat positions 1, 3, and

4. Available inputs were limited to exterior microphones 3 and 4, and accelerometers

5Y and 5Z. Additional available transducers included accelerometers 1 and 2, plus four

accelerometers mounted on the pressure bulkhead (which were not available for the 8x8

analysis; see Figure 3-5 for locations).

Table 8-3 lists the ordinary coherence values between the various inputs and out-

puts, for BPF(8) and BPF(10) tones. The table shows that there is a highly coherent

environment on the Demonstrator for the 10x8 engine configuration. The coherence

values are all nearly unity, except for coherence values of about 0.7 between accelerom-

eter 5Z and each output microphone for the BPF(10) tone. These uniformly high

coherence values are in contrast to the coherence values for the 8x8 data (Table 8-1),

where the two sources of the same frequency combined to generate incoherent signals.

Conditioning the three output microphones with the various inputs results in the

reductions in BPF tone levels listed in Table 8-4. As for the 8x8 data, the reductions

in levels shown in the table due to conditioning with the individual inputs reflect the

coherence values. The results of conditioning with the combined inputs suggest that

both acoustic and vibration excitation contribute to the levels at the output micro-

phones, at about equal strength. However, since there is high coherence among the

various inputs for both BPF tones, the relative importance of acoustic and vibration

sources to the outputs cannot be determined with certainty.

Conditioning accelerometer 2, the bulkhead accelerometers, and aft section micro-

119

phone 5 as outputs with the same inputs as above also indicates approximately equal

contributions from acoustic and vibration sources for each of these outputs. Again , since

the levels between these output transducers are all highly correlated with each other

as well as with all the available input transducers, and since several input transduc-

ers are missing for this 10x8 analysis, no further conclusions can be drawn concerning

transmission paths from the acoustic and vibration sources into the passenger cabin.

120

TABLE 8-1CoherenceBetweenInputsandOutputs

for BPF Tone(168Hz}, Testpoint19550A03

Inp_tTransducer

Mic 4 0.30 0.38 0.98 0.97 0.40

13 0.08 0.75 0.75 0.77 0.79

3 0.80 0.18 0.78 0.65 0.08

12 0.84 0.50 0.45 0.28 0.19

Accel 4 0.79 0.05 0.34 0.4_6 0.47

5Z 0.70 0.31 0.78 0.60 0.03

5X 0.60 0.39 0.87 0.69 0.03

5Y 0.50 0.39 0.94 0.81 0.18

121

TABLE 8-2

Reduction in BPF Tone Level at

Interior Output Microphones Conditioned with

Exterior Microphone and Accelerometer Inputs,Test Point 19550A03

Input _eTransducer 6

Mic 4 1.5 2.0 17.5 15.0 3.0

13 0.0 6.0 5.5 6.0 7.0

3 7.0 1.0 6.0 4.5 0.5

12 8.0 3.0 2.5 1.0 0.5

Mic 4 1.5 2.0 17.5 15.0 3.0

+13 15.0 11.5 20.5 21.5 7.0

t- 3 15.0 12.0 21.0 21.5 7.5

+12 15.0 12.0 21.0 21.5 8.0

Accel 4 6.0 1.0 1.5 2.5 3.0

5Z 6.0 2.1) 6.5

5x :;.0 2.0 8.5

5Y 3.0 2.0 12.0

Accel 5Z 6.0 2.0 6.5

+5X 9.0 6.0 15.5

+5Y 12.5 10.0 19.5

+4 13.0 11.0 22.5

4.0 0.5

5.0 0.5

7.5 1.0

4.0 0.5

12.5 7.0

16.5 12.0

18.5 13.0

Accel 5Z

+5X

_-5Y

F4

+Mic 4

+13

t-3

....... +12

6.0

9.0

12.5

13.0

13.0

15.5

15.5

15.5

2.0 6.5 4.0 0.5

60 115.5t12.51 7010.0119.5116.5 [ 12.0

11.0t22.5 t 18.5 I 13.0

11.5 23.5 22.0 ] 15.5

13.0 / 24'5 / 22"5 I 18.5

I I 19.0( '_ 5 ')') 5 q14 )[ .4, _. I 0

122

TABLE 8-3

Coherence Between Inputs and Outputs

for BPF Tone (168 and 210 Hz), Test Point 19640B03

Input Output Microphone

Transducer 1 3 4

BPF(8) BPF(10) BPF(8) BPF(10) BPF(8) BPF(10)Mic 4

3

Accel 5Z

5Y

0.98

0.98

0.98

0.96

0.94

0.93

0.72

0.95

0.95

0.95

0.95

0.93

0.92

0.89

0.69

0.92

0.97

0.97

0.98

0.95

0.97

0.96

0.75

0.97

123

'FABLE 8-4

Reduction in BPF Tone Level at

Interior Output Microphones Conditioned with

Exterior Microphone and Accelerometer inputs,Test Point 19640B03

Input Output Microphone

Transducer 1 3 4

Mic 4BPF(8)

16.5

16.0

L ic,II165+3 16.5

BPF(10)

I 17.514.0

12.5

11.5

12.5

13.0

BPF(8)13.0

13.0

13.0

13.5

BPF(10)11.0

9.5

11oII11.5

Accel 5Z 5.5 13.0 5.0

5Y 12.5 Ii.0 II.0

5.5

12.5

13.0

13.5

13.0

13.5

13.5

14.0

5.5

12.5

13.0

13.5

Accel 5Z [[ 17.5+SY 17.5

Accel 5Z

+5Y

+Mic 4

+3

17.5

17.5

18.0

18.5

5.0

11.0

5.0

11.0

11.5

12.0

BPF(8) BPF(10)13.5 14.5

13.5 13.5

13.5 14.5

14.0 15.0

14.5 6.0

12.5 14.5

14.5 6.0

15.0 15.0

14.5 6.0

15.0 15.0

15.0 15.5

15.0 15.5

124

I7g848175

81848378

(JT8D Side) EngineP_lonBulkhead

Right,Sidewall

AftBulkheads

(-60)

85 7083 8S

81 8584 77

Aft Pressur_ Engine Pylon

Bulkhead _ Bulkhead.

-83 87 91 76-78 85 84 77

66 81 80 83-69 73 79 74

(UDF Side)

LP,ft:

Sidewat[

-76 -7977 -7581 -79

-65 -73

-78 84 84 76 (-62) -75-83 82 80

(-62) 79 -82-70 82 77 76 79 -79-67 73 72 65 67 -71

ONE-THIRD OCTAVE BAND

CENTER FREQUENCY

IN EACH BOXED AREA

First Row, JNI: I00 Hz

Second Row, BPF: 160 HzThird Row, UNit 200 HZ

Fourth Row, 2BPFt 315 HI72 66 -71

-75 78 78 -82 (-61) -7873 79 79 -78 -66 -71

-73 80 76 63 (59) -68-71 75 73

-77 77 79-83 80 78-78 77 76-68 70 73

-69 -67 -66-84 -79 -77-75 -68 -75

68 -62 -69

Lower Upper Oag

Sidewall SEdewatl Back

Bag Upper LowerBad< SidewallSidewall

-79-82-78-65

-74(-?1)

-7669

-75-80-78

73

-78-82-74

72

Fvd Aft

64-71

(-66)71I-;;-71

I .... J

72 7784 86

79 Floor69 77

FIGURE 8-1. Measured Sound Intensities for the 100, 160, 200 and 315 Hz

One-Third Octave Bands, 8x8 Flight 2026.

125

t 77

6280

_ 744

81848278

(JT8D Side)

ftiyhLSidewall

'i"'--i iill-66

-76-80-75-64

I-72 76 76 I

70 77 77 ]-70 75 74 I-68 73 71 ]

Aft l/ulHleads

....... t ........ !Ii

-74 I 75 77-St ! 78 76-75 =i 75 74-70 66 71I

l

Hack Sidewall 5ide*_l[

En9inePyLonBuLkhead

(-60)92eS83

848283

BuLkhead _

76 \8)

"/75

70828472

ONE-THIRD OCTAVE BAND ]CENTER FREQUENCY ]

IN EACH BOXED ARIA ]!

Ficst Eow, JNI= 100 Hz [Second Row, BPF= 160 B= |

Third Row, UNI: 200 HI |Fourth ROW, 2BPF= 315 R=I

I

(UDF Side)

74 -7475 7581 7972 -63

-,,[-72-76

74 (-60) -731(-SO) 77 -79

74 77 -76163 65 -68

I

.....................70 6_ -6_

-O0 1-591 -75-76 -64 -691

61 (57) 65[

.............. 6il-64 -62-79 -74 -72-70 -63 -701

63 -58 -631

.................... i

tnvcr I_p_ 8_SidcvaLL SidevalL fla,;k

-78-81-77-64

-73(-69)-75

68

-73-79-77

73

-77-81-73

71

Fwd AlL

- 63-71

72 7584 04

(63) Flnor68

FIGURE 8-2. Measured Sound Power Levels for the 100, 160, 200 and 315 Hz

One-Third Octave Bands, 8x8 Flight 2026.

126

o

o--

E--

L_

o

2_

g_

BPF

A/,,AAA

50 I i]O 150 200 250 300

Frequency4 IHz}

VIEW A-A

350 400 450 500

STA. 1271

200 250 300 350 '100 450 500

Frequencq IHzl

A

SEAT ROW 6

SEAT ROW 4

^

Ion Root

200 250 300 350 400 450 500

Frequenc q [Hz)

FIGURE 8-3a. Cabin Sidewall Transmission Path: Exterior Noise

Spectra Measured on the Aircraft Fuselage.

Test Point 19550A03.

127

0

°,

BPFJN1 UN1• i i I ,

; !i

--' ' " t I i

Jl-J,i i

VIEW A-A

2BPFI .

| i

| w

.......... I .... I ............

200 ZSO :300 3SO 400 450 500

Frequent _i IHzl

u6_

_o ........ , ............. , ....SO I O0 l SO _00 250 300

Frequenc q {Hz)

350 4O0 450

//_ Planes of

STA. 1271 /

i -r v .,

..................._',

('1,(; .:_:,.3 I ....... 1"_'- !

'1_--- RP

I

I

_A _ Pylon Root

__. "-- SEAT ROW 6SEAT ROW 4

rotation

FIGURE 8-3b. Cabin Sidewall Transnfission Path: Interior Noise

Spectra, Test Point 19550A03.

128

0

no0

Loo

E--

CD , ,

_= IIIt.@

0

I

I

//t

l_,l I_ l

" 11 \IV

B]'F

SI II

/1

I '\1I

.... i ............................

50 I00 150 200 250 300 350 400

Frequency (Hz]

£C3

0..

0L

.3E

0CxI

.$L

oo

l

\\

\\\

_R

I• - i ....

45O 5OO

'F

i

\

2BI

o so ,oo ,so aoo '2so' "_oo'"Frequenc_l IHz)

i I

I II I

i I, i

, !

l i

FIGURE 8-4. Aft Section Transmission Path: Unpressurized Section

Noise Spectra, Test Point 19550A03.

129

cn_-"0

gl

g-o

a_g-"0

(j?l

g-o

0

IL 0

0

L0

tDIM C-_

.°

%e-

N-o

EL i iTest Configuration 4: 'Quiet Cabin' ---

2OO 400

t I 1 1 I 1 ICon_:.r.tion6:,,,See,on,bso.p.o..en,o,edI--4'

.---BP

200 400 600 800 1000 1200 1400 1600 1800 2000

FFequenc_j (Hz)

I I I 1____21 J? "lYst Configuratirm 7" Double Wall Bulkhead Ite,noved

_Jv"_-'_ ,-L-..... I----It---SBPF-gBPF 10BPF-__

200 400 600 800 1000 1200 1400 1600 1800 2000

Frequenc q (Hz]

FIGURE 8-5. Aft Section Transmission Path: Comparison of Interior

Noise Spectra at Row 6, Seat 1.

130

t_

o

f-'x

m

O

co

e

O

,(L..

o

Q_o3

o

o

1160

Exterior Fuselage Microphones _

0

Fuselage Accelerometer

h|terior Unpressurized Section __

1100 1200 1220 !240 1260 1280 1300

Rotor Speed (rpm)

FIGURE 8-6. Comparison of Noise and Vibration Data Acquired in the

Aft Section of the Aircraft, Flight 1955.

131

Exterior Mics

f

j/

Pylon/Fuselage Accels

Interior Mics

Row 6 Mics

Q.q2

FIGURE 8-7. Transducers Used in the Partial Coherence Analysis

of 8x_ Data

132

9 Summary and Conclusions

Noise and vibration data acquired during flight tests of the UHB Demonstrator with

a GE UDF engine were analyzed to evaluate the effectiveness of selected noise control

treatments in reducing interior noise levels, and to investigate cabin noise characteristics

and transmission paths. Most of the flight data were acquired at high altitude, high

speed cruise conditions (35,000 ft. and 0.76 Mach), with an 8x8 engine configuration.

Additional tests were conducted with a 10x8 engine configuration and at lower altitudes

and speeds, permitting a comparison of interior noise levels among the various engine

configurations and operating conditions.

Noise control treatments were installed on the Demonstrator to reduce cabin noise

to a maximum level of 82 dBA. Seven configurations were evaluated; in the first four,

treatments were added up to the Quiet Cabin configuration for which the goal level

of 82 dBA was achieved. The final three configurations involved removal of selected

treatments without increasing cabin noise levels. The treatment evaluation analysisshowed that:

• the torque box damping and the engine dynamic absorbers have a negligible effect

on cabin noise levels,

• the cabin furnishings (including damped trim panels) reduce cabin noise levels

significantly, and

• removal of engine dynamic absorbers, cargo skin damping, and double wall pres-

sure bulkhead treatments has a negligible effect on cabin noise levels.

Interior noise spectra in the aft cabin were found to contain several tones, super-

imposed on a broadband background. These tones correspond to the blade passage

frequencies of the two propellers, the propeller shaft rotational frequency, the core ro-

tational frequencies of the UHB and JT8D engines, and various harmonics of these

several tones. The broadband background is due to the exterior boundary layer noise.

In terms of A-weighted sound levels, at a particular seat location the broadband

component of the noise level is typically equal to or higher than the tonal component

(which is the summation of the levels of all the tones in the spectrum). Since there is

little variation in the broadband noise in the aft cabin, the range in overall noise levels

is small, typically from 78 to 82 dBA over the last five seat rows.

The main contributors to the tonal component, in order of importance, are the BPF

tones, the JN1 tone, the UN1 tone, and the 2BPF tones. The BPF tone levels increase

with increasing rotor speed, reflecting the increase in exterior noise levels observed in

133

the forward rotor plane as rotor speed increases. The BPF levels are typically highest

in the aft cabin, oil the right side of the aircraft.

Comparison of interior noise levels for high speed, high altitude cruise operations

measured during 8x8 and 10x8 flights showed comparable maxinmm overall levels for

the two engine configurations. Broadband levels were nearly identical; BPF tone levels

were lower at selected locations during the 10x8 flights. These lower levels correspond

to the measured lower noise levels in the forward rotor plane and lower vibration levels

on the engine mount for the 10x8 flights.

Noise data from several flights at lower altitudes and speeds than the standard

35,000 ft., Mach 0.76 cruise conditions were reviewed for 8x8 and 10x8 configurations.

For these "non-cruise" conditions, the levels of the broadband and tonal components

were different than for cruise conditions. Differences in the broadband levels were

traced to expected differences in exterior boundary layer noise levels (due to altitude

and speed differences). Differences in the tonal levels corresponded in most cases with

differences in exterior noise levels in the forward propeller plane, which are attributable

to altitude and helical tip Mach number effects.

The sound intensity survey showed that most of the tone energy in the cabin em-

anates from the engine mount bulkheads and the pressure bulkhead (if untreated), and

tile sidewall on the JT8D side. Study of the various interior and exterior noise and

vibration data, in conjunction with the results of the sound intensity survey and the

partial coherence analyses, yields the following conclusions about noise sources and

transmission paths into the cabin:

the cabin sidewall path is not a propeller noise path but is the likely path of the

broadband noise component,

the aft section path does not appear to be a transmission path for the propeller

noise at low frequencies, but if left untreated may allow higher harmonics of the

propeller noise to be transmitted and become perceptible in the cabin, and

the tones found in the interior noise spectra are the result of structureborne

transmission through the pylon and the fuselage. The transmitted energy is

caused by a combination of acoustic loads from the propeller blades impinging

on the pylon, and mechanical forces within the engine propagating through the

pylon. The major radiating surfaces are the aft bulkheads (primarily the engine

mount bulkheads) and the right (JT8D side) sidewall.

Finally, two analysis techniques were evaluated using measured Demonstrator flight

test data. The partial coherence analysis technique was found to be very useful in

identifying the relative strengths of the sources of BPF tones in the cabin, but of limited

134

usefulness in sorting out transmission path strengths. The weakness of this technique

was in part related to the limited availability of transducers, at appropriate locations.

Nevertheless, the conclusions drawn from the partial coherence analyses agreed with

conclusions developed independently from other data analyses, demonstrating that the

technique has good potential for source/path diagnosis.

The second technique (described in Appendix A) is a spectral analysis method

of estimating turbulent boundary layer pressure fluctuations using a fixed microphone

probe on the fuselage exterior. Good agreement was found between predicted levels and

wavenumber-frequency levels estimated from the flight data, in the mid-wavenumber

and frequency range. This demonstrates that the fixed probe technique works, and can

be used successfully to measure turbulent boundary layer spectra in flight.

135

A Estimation of The Turbulent Boundary Layer

Pressure Wavenumber-Frequency Spectrum

The pressure and vibrational fields associated with turbulent flows over various

vehicles have been the subject of continuing research over the past 30 years. In com-

mercial aircraft, the turbulent boundary layer (TBL) pressure fluctuations represent

one of the dominant sources of interior noise during cruise, as was also observed on

the UHB Demonstrator. The modeling of the turbulent boundary layer pressure fre-

quency spectrum for interior noise prediction is generally based on an empirical fit to

laboratory wind tunnel measurements. As a result, cabin flight test data are not in

good agreement with predicted levels based on such models. In addition, it has been

reasoned that the absence of wavenumber information leads to an overprediction of

interior levels. Wavenumber information is also useful since the relative importance of

nonresonant transmission at low wavenumber can be determined by the wavenumber-

frequency spectra of the TBL fluctuations.

The main objective of the study described in this Appendix is to examine a new spec-

tral analysis method of estimating turbulent boundary layer pressure fluctuations using

a fixed probe consisting of two microphones. Demonstrator flight test data have been

used to examine the applicability of this spectral analysis method to in situ flight mea-

surements of TBL pressure fluctuations. TBL pressure wavenumber-frequency spectra

on the fuselage surface were determined from measurements using two exterior mi-

crophones, for several flight conditions. These estimated inflight spectra were then

compared with existing wind tunnel data, and with selected prediction models. The

behavior of the TBL wavenumber-frequency spectra in the lower wavenumber region

was also examined with regard to the "wavelength white" region of the wind tunnel

TBL spectra.

A.1 Definition of Wavenumber-Frequency Spectra

Turbulent flow is a random process in both space and time, and cannot be specified

deterministically. The turbulent pressure field is, therefore, modeled as a zero-mean

random process with a space-time correlation function given by

Rvv(rl, r2; tl, t_) = £{p(rl, _l)p(r2, t2)},

where rl and r2 denote two observation points, and £ is the expectation operator. A

comnmn property of the random field is temporal stationarity which implies a depen-

dence of Rpv only on the difference tl - t2, i.e.

136

R_(rl,r2;tl,t_)= Rpp(rl,r2;tl- t2,0),

= R_(rl,r2;_-),

where "r- tl- t2.In t.hiscase the fieldcan be characterized in terms of itscross-spectral

density as follows:

Spp(rl, r2; w) / Rpp(rl,---- r2," r) e-i_r dr,

which is obtained as the Fourier transform of R_, over the r variable.

Another property of some fields is spatial homogeneity which implies that R_ de-

pends only on the spatial separation rl - r2, giving:

Rpp(rl,r2;T)---- R,,p(r;'r)

and

where r -- ri - r2. The wavenumber-frequency spectrum can now be defined as the

Fourier transform of the spatial-time correlation on the spatial separation vector r and

the time difference v. For spatially homogeneous fields, the wavenumber-frequency

spectrum is given by:

S_(k, w) =//Rpp(r; r) e -i(k'r+wr) drdT.

The wavevector k is the Fourier conjugate variable of the spatial separation vector r,

and the circular frequency w is the conjugate variable of the time difference v.

For a nonhomogeneous field, the wavenumber-frequency spectrum can be defined

as a space-varying spectrum, i.e.,

S_(x, k, w) = ]] Rpp(x, r; r) e-i(k.r+_r) dr dr.

The space-averaged spectrum for a nonhomogeneous process may be written:

1/S_(k,w;A) = _ S_,(x,k,w)dx,

where A is the area of the measurement surface.

The measurement of the wavenumber-frequency spectral density, in general, requires

an array of transducers carefully placed in the field under investigation. The space-

time sampled data is then Fourier transformed to obtain the wavenumber-frequency

spectrum.

137

A.2 Spectral Modeling of TBL Pressure Fluctuations

The pressure fluctuations in an incompressible flow have classically been examined

by relating the phenomenon to velocity fluctuations through Poisson's equation. Some

of the theoretical knowledge seems to be contradicted by experiments, however, par-

ticularly the behavior of the spectrum at long wavelengths. For example, the theories

of Kraichnan [1,2] and Phillips [3] indicated that the spectrum should vanish at long

wavelengths whereas the experimental results are best described by models in which

the spectrum approaches a constant in that region, e.g. Corcos' model [4]. As a result,

several senli-empirical/empirical models have been developed primarily on the basis of

experimental data. (A complete description of the vast amount of work related to TBL

theoretical and empirical models is beyond the scope of this report.)

In the present investigation, the TBL wavenumber-frequency spectra obtained from

flight test data were compared with three different prediction models: Corcos' convec-

tive ridge model [4], a spectral model proposed by Laganelli et al. [6], and Witting's

burst and sweep model [5]. These models appear to match a broad variety of experi-

ments and provide a basis for wavelength-white spectra at long wavelengths.

Many measurements of two point wall pressure statistics were carried out through

the 1960's and early 1970's. Those conducted using cross-spectral techniques, e.g. Bull

[7] and Blake [8], are of more significance for TBL pressure spectrum estimation than

those conducted using cross-correlation techniques, e.g. Willmarth and Wooldridge [9].

The cross-correlation measurements did not provide sufficient accuracy when Fourier

transformed to the frequency ranges of interest for such problems as aircraft cabin

noise.

On the basis of the wind tunnel measurements, it has been recognized that the

wavenumber-frequency spectrum of the wall pressure is sharply peaked with regard

to the longitudinal (streamwise) wavenumber at a value kl _ w/Uc, where Uc is the

convection velocity and w is the circular frequency. The decrease with lateral (span-

wise) wavenumber /¢3 is much slower, so that S_(k,w) posseses a convective ridge

centered at (kl, k3) _ (w/Uc, 0) and is oriented towards the flow. Wind tunnel mea-

surements by Willmarth and Wooldridge [9] and Corcos [4] support the conclusion that

Spp(ks) _ Sin,(0). The fundamental problem, then, is the description of the longitudinal

component Sin, (kl).

Although a large number of laboratory wind tunnel measurements have been per-

formed, only a few have addressed the wavenumber-frequency description, particularly

in the lower wavenumber-frequency region of the turbulent fields. Blake and Chase

[10], Farabee and Geib [11], and Martini, Leehy, and Moeller [12] among others used a

flush mounted array of microphones to measure the low wavenumber turbulent pressure

spectrum. Wavenumber ineasurements were also carried out using structural elements

138

such as membranes and plates, e.g. Martin and Leehy [13] and Jameson [14].

Figure A-1 shows a possible shape of the wavenumber spectrum for zero lateral

wavenumber and at a fixed frequency in terms of the longitudinal wavenumber [5,12].

The shape of this spectrum near the convective ridge ¢o/Uc is rather well established

by cross-spectral measurements, but the shape near and below the acoustic wavenum-

ber, ka = _v/c0 (where co is the speed of sound in the ambient medium), has not been

conclusively established. The figure also shows two possible spectral characteristics in

the low wavenumber region: the upper dashed curve shows a flat region where S_ is

independent of kl, while the lower solid curve has a region where Spp is proportional

to kl2, and so would tend to zero in the absence of the acoustic peak. Most exist-

ing low wavenumber wall pressure measurements become independent of wavenumber

("wavenumber white") beginning at a wavenumber substantially above the acoustic

wavenumber, and thus favor the upper curve. There is no indication that the incom-

pressible k_ low wavenumber limit of Kraichnan [1], as shown in the lower curve, is

approached in any way.

On the other hand, an extensive search of the available published literature has

revealed that no attempt has been made to estimate, in situ, the wavenumber-frequency

spectrum of the turbulent boundary layer pressure spectrum on the exterior of all

airplane fuselage. Most of the flight test investigations reported in the past are limited

to pressure power spectral density, correlation, and cross spectrum analysis [15-17].

The estimation procedure used here for determining the wavenumber-frequency

spectrum is unique in the sense that it requires the use of a fixed probe consisting of

two microphones and can be easily applied to inflight measurements. The theory and

assumptions behind this estimation technique are described in the following sections.

A.3 Estimation Procedure

As mentioned earlier, wavenumber-frequency spectra can be determined through

measurements using an array of carefully placed transducers. In recent years, new

techniques applicable to homogeneous and stationary fields have been developed for

estimating wavenumber-frequency spectra which use two fixed transducers for data

acquisition. These new methods have been applied to studies of ion-acoustic and plasma

turbulence [18-21].

In one method developed by Harker and Ilie [18-19], the spatial Fourier transform is

performed on the covariance of pressure signals measured by a movable probe at many

sequential values of probe separation. In many cases, however, it is impractical to vary

probe separation to obtain information over a sufficiently wide range of wavenumber.

In order to avoid this difficulty, Iwama, Ohba, and Tsukishima [20] developed a cot-

139

relation method of measuring the first and second moments of the spectral density

wavenumber space using fixed probe pairs based on the fact that

/ k%(k]w)dk,i,qt(0, ) _

where H(_,w) is the cross-spectral density from the two transducers (separated by a

distance _):

1

H(_,w) : _ f S(k,w)e ''_dk,

and s(kl,_ ) is the conditional wavenumber spectral density:

S(_o) is the auto spectral density defined by:

/.

= = ] dk.

Introducing the concept of local wavenumber-frequency spectral density, Beall, Kim,

and Powers [21] developed a new approach for estimating wavenumber-frequency spec-

tra using fixed probe pairs. The local wavenumber is analogous to instantaneous fre-

quency in the time domain. This local wavenumber method was used in the present

investigation for estimating wavenumber-frequency spectra of the TBL pressure fluctu-

ations from Demonstrator flight test data. The details of the theoretical approach to

this method are given in Reference 21.

Most turbulence theories model tile fluctuations in a turbulent medium as a homo-

geneous and stationary process [1-3]. Kraichnan's analysis of anisotropic homogeneous

turbulence is cast mainly in terms of the Fourier coefficients of the velocity field. The

turbulent field is supposed to be spatially periodic with very large period L, and the

velocity may then be expressed as a Fourier series:

u(x,t) = _ A(k,t)e ik'x,

where the summation is over all the wavenumbers permitted by the cyclic boundary

conditions. This analysis permits the representation of the fluctuations in a turbulent

medium by a superposition of wave packets [21] or oscillations, which are approximately

sinusoidal in both space and time with slowly varying amplitudes and wavenumbers.

Under such conditions, Beall et al. showed that the local wavenumber-frequency spec-

tral density is the same as the conventional wavenumber and frequency density [21].

140

In addition, a basic of property of wavefieldsthat

Ox '

just as in the time domain the circular frequency

00(t, )03 --

Of '

has been ingeniously used to obtain the local wavenumber.

The local wavenumber can be readily estimated from two fixed probe pairs at ah

and x2:

X 2 -- X 1

where

X 1 -_- X 2X--

2

The spectral densities H(0,_), H(_,w), and St(k,w) may be estimated by using two

probes at x = xl and x2 separated by a distance _ = xl - x2. Here, St(k,w) is the local

wavenumber -frequency spectral density.

For a zero-mean, stationary, homogeneous random field ¢(z, t), the Fourier trans-

form relationship

O(z,w) = f ¢(x,t)e -i_'t dt,

can be represented by the following discrete Fourier series:

1

=

The sample cross spectrum is:

N

¢(z,eAt)e -'_latl=l

H(JX_,w) = ¢(_)(xl,w)¢(J_x2,w)

= +

where C (j) and Q(J) are, respectively, in-phase and quadrature components of the cross

spectrum. The local wavenumber is given by:

'

141

where

00)= tan-1 Q(J_)

The estimated wavenumbers are restricted to the interval -Tr/_ to 7r/_ if indetermina-

cies of -i-n21r/_ are to be avoided.

The local wavenumber and frequency spectrum St(k,_) is computed by summing

the sample power values S(J_o_) at a fixed frequency from those records which have a

sample local wavenumber in the range k to k + Ak, and dividing by M:

1 M

S,(k,_v) = _ _ I[0,ak)(k - k(J_o.')) s(J{_).j---1

The indicator function I[0.h)(x) is defined as

1, O__x<hIi0,h)(x) = O, otherwise

A.4 Flight Test Results

The wavenumber-frequency spectra for five flight conditions of the Demonstrator

(see Table A-l) were estimated using the local wavenumber approach. The second

flight condition is very similar to the first and was chosen to check the repeatability

of the local wavenurnber method. The locations of the probe microphones, exterior

microphones 9 and 10, are shown in Figure 3-1 in the main report text. The spacing

between these two microphones is 17.75 inches, which permits an estimation of the

wavenumber spectra in the range of -2.1 to 2.1 ft -1.

The sound pressure spectra with the two probe microphones for all flight condi-

tions are shown in Figures A-2a through A-2e. It may be observed that the difference

in overall sound pressure level between the two microphones for the first two flight

conditions is about 5 to 6 dB, while for the other cases it is about 2 to 3 dB. The

three-dimensional plots of the estimated wavenumber-frequency spectra for the TBL

pressure fluctuations are correspondingly shown in Figures A-3a through A-3e, while

Figures A-4a through A-4e show the contour plots of the pressure fluctuations. The

variation of the TBL pressure spectra, and in particular the variation of the convective

ridge peak with wavenumber and frequency, may be seen in these figures.

In order to compare the flight test data with the available wind tunnel or prediction

models, the TBL pressure spectral density was normalized with respect to the following

flow parameters: the free stream dynamic pressure (q_), the boundary layer displace-

ment thickness (_f'), and the free stream flow velocity (U_). The wavenumber k and the

142

circular frequency w were non-dimensionalized as k6" and w6*/Uoo (Strouhal Number),

respectively. Wind tunnel measurements give convective velocity Uc typically in the

range of 0.6 to 0.8Uoo, with the larger Uc corresponding to longer wavelength scales.

Inflight measurements by Bhat [15] have shown Uc to approach an asymptotic value of

0.78Uoo with increasing microphone separation. A value of 0.76Uoo was chosen for Uc

in our analysis.

Figures A-5a though A-5d show the variation of pressure spectral density in the

wavenumber plane for fixed Strouhal numbers. These TBL pressure plots exhibit a

pronounced peak occurring at or near the corresponding convective wavenumbers given

by kc = w/U_. The location of the convective peak agrees well with that predicted by

theoretical considerations. Since the acoustic and convective wavenumbers (ka and

kc, respectively) are fairly close together, the TBL pressure spectrum between them

is dominated by the convective ridge pressures. Above k_, the TBL pressure decays

rapidly.

The estimated TBL pressure wavenumber-frequency spectra were compared with

three prediction models: Corcos' convective ridge, Laganelli's, and Witting's burst and

sweep. The Corcos convective ridge model [4,13], based on similarity laws and slightly

modified using Blake's cross-spectral density data [8], was used for comparison with

the current flight test data. Figure A-6 shows a comparison of the estimated TBL pres-

sure wavenumber spectra with those obtained from Corcos' model. The current TBL

spectral levels are found to be much lower than those calculated from the convective

ridge model. Martin and Leehy [13] have reasoned that due to the breakdown of sim-

ilarity laws in the lower wavenumber region, the convective ridge model considerably

overestimates the TBL spectral levels in this wavenumber range.

The estimated TBL spectral levels are compared in Figure A-7 with those obtained

from Laganelli's model [6]. The figure shows that the flight test data for conditions

4 and 5 are in very good agreement with the values give by the prediction model,

while data for conditions 1, 2, and 3 show a deviation of 4 to 7 dB. The reasons for

this discrepancy with predictions are difficult to assess with the limited amount of

data available. One possible explanation may be found in the location of the two-

micropohone probe, which was upstream from the modified pylon supporting the UHB

engine. Flight test data has shown that there is a rapid acceleration of local airflow

just upstream of this pylon. The turbulent boundary layer displacement thickness

was found to be approximately 20% larger than that observed further upstream of the

microphone probe. This may result in higher pressures at the microphone locations.

Figure A-8 shows a comparison of the normalized (convective ridge peak) wavenum-

ber spectral density with the peak values of Witting's model spectra [5]. The normal-

ization constant chosen for Witting's model is based on the convective velocity, U_, and

the overall sound pressure level. The estimated spectral levels are seen to be higher

143

(about 10 to 12 dB) than those obtained from Witting's model, particularly at very low

wavenumbers. Several researchers in the past have observed that wavenumber spectral

measurements in the low wavenumber range are very difficult to make and are subject

to error. Further, there is considerable scatter among the various wind tunnel data in

the low wavenumber region. It may also be mentioned that Witting's model is valid

for incompressible flow velocities (Uc << co).

In summary, the wavenumber-frequency TBL pressure spectra estimated from flight

test measurements were found to be on the high side when compared with Laganelli's

and Witting's prediction models. Overall, the measured sound pressure levels, the

normalized TBL pressure, and the wavenumber spectra are consistently higher than

the predicted values, particularly in the very low wavenumber region, and might be

an indication of the flight turbulent activity. The estimated wavenumber-frequency

levels are in very good agreement with the predicted data in the mid-wavenumber and

frequency range. The results presented in this appendix, therefore, show that the local

wavenumber technique using a fixed probe can be used for estimating the wavenumber-

frequency spectra of the TBL pressure fluctuations. This estimation method can be

easily used to measure in situ TBL pressure wavenumber-frequency spectra during

flight.

Further tests using the microphone array method should be conducted to exam-

ine the underlying assumption of homogeniety of the TBL pressure field. Additional

measurements with smaller probe nficrophone spacing should also be undertaken, to

expand the range of wavenumber spectra.

o

°

.

.

References

Kraichnan, R.H., "Pressure Field Within Homogeneous Anisotropic Turbulence,"

J. Acoust. Soc. Am., 28(1), 64-72 (1956).

Kraichnan, R.H., "Noise Transmission from Boundary Layer Pressure Fluctua-

tions," J. Acoust. Soc. Am., 29(1), 65-80 (1957).

Phillips, O.M., "On the Aerodynamic Surface Sound from a Plane Turbulent

Boundary Layer," Proc. Royal Soc. Lond., A234, 327-335 (1956).

Corcos, G.M., "The Structure of the Turbulent Pressure Field in Boundary Layer

Flows," J. Fluid Mech., 18, 353-378 (1964).

Witting, J.M., "A Spectral Model of Pressure Fluctuations at a Rigid Wall

Bounding an Incompressible Fluid, Based on Turbulent Structures in the Bound-

ary Layer," Noise Control Eng. Journal, 26(1), 28-43 (1986).

144

6. LaganeUi,A.L., MarteUucci, A. and Shaw, L.L., "Wall Pressure Fluctuations in

Attached Boundary-Layer Flow," AIAA Journal, 21(4), 495-502 (1983).

7. Bull, M.K., "Wall Pressure Fluctuations Associated with Subsonic Turbuluent

Boundary Layer Flow," J. Fluid Mech., 28(4), 719-754 (1967).

8. Blake, W.K., "Turbulent Boundary Layer Wall Pressure Fluctuations on Smooth

and Rough Walls," J. Fluid Mech., 44(4), 637-660 (1970).

,

10.

11.

12.

13.

14.

15.

16.

17.

WiUmarth, W.W. and Wooldridge, C.E., "Measurements of the Fluctuating Pres-

sure at the Wall Beneath a Thick Turbulent Boundary Layer," J. Fluid Mech.,

14(2), 187-210 (1962).

Blake, W.K. and Chase, D.M., "Wavenumber-Frequency Spectra of Turbulent

Boundary Layer Pressure Measured by Microphone Arrays," J. Acoust. Soc.

Am., 42(3), 862-877 (1971).

Farabee, T.M. and Geib, F.E., "Measurement of Boundary Layer Pressure Fields

with an Array of Pressure Transducers in a Subsonic Flow," David Taylor Naval

Ship Res. and Dev. Ctr. Report No. 76-0031 (1976).

Martini, K., Leehy, P. and Moeller, M., "Comparison of Techniques to Measure

the Low Wavenumber Spectrum of a Turbulent Boundary Layer," Acoustics and

Vibration Laboratory, MIT Report No. 92828-1 (1984).

Martin, N.C. and Leehy, P., "Low Wavenumber Wall Pressure Measurements

Using a Rectangular Membrane as a Spatial Filter," J. Sound Vib., 52(1), 95-120

(1977).

Jameson, P.W., "Measurements of the Low Wavenumber Component of Turbu-

lent Boundary Layer Wall Pressure Spectrum," Bolt Beranek and Newman Inc.

Report No. 1937 (1970).

Bhat, W.V., "Flight Test Measurements of Exterior Turbulent Boundary Layer

Pressure Fluctuations on Boeing Model 737 Airplane," J. Sound Vib., 14(4),

439-457 (1971).

MaestreUo, L., "Boundary Layer Pressure Fluctuations on the 707 Prototype

Airplane," Paper presented at the 64th Meeting of the Acoustical Society of

America, Seattle (1962).

Heron, H.K. and Webb, D.R.B., "Flight Measurements of Pressure Fluctua-

tions in a Turbulent Subsonic Boundary Layer and the Relation with Wall Shear

Stress," R.A.E. Tech. Report No. 66338 (1966).

145

18. Harker, K.J. and Ih_:, D.B., "Measurement of Plasma Wave Spectral Density

from the Cross-Power Density Spectrum," Rev. Sci. Instrum., 45(11), 1315-1324

(1974).

19. lli¢', D.B. a,ld llarker, K.,I., "F, vahlati.,i ,_f I_htsma Wave Spectral I}ensil.y fr_mh

Cross-Power Spectra," Rev. Sci. lnstrum., 46(9), 1197-1200, (1975).

20. Iwama, N., Ohba, Y. and Tsukishima, T., "Estimation of Wavenumber Spectrum

Parameters from Fixed Probe Pair Data," J. Appl. Phys., 50(5), 3197-3206

(1979).

21. Beall, J.M., Ki,n, Y.C. and Powers, E.J., "Estimation of Wavenumber and Fre-

quency Spectra Using Fixed Probe Pairs," J. Appl. Phys., 53(6), 3933-3940

(1982).

22. Corcos, G.M., "Resolution of Pressure in Turbulence," J. Acoust. Soc. Am.,

35(2), 192-199 (1963).

146

Table A-1. Flight Test Conditions and Parameters

Flt.

Cond.

No. Fit/Point Mach No.

Pressure True

Altitude Airspeed

(St) (knots)

Boundary Layer

Displacement

Thickness

_', (inches)

Dynamic

Pressure

q®,(Pa1

2

3

4

5

19530A02 0.76

19520A05 0.76

20330K04 0.77

19470002 0.57

19480004 0.81

35000 444

35000 445

22600 472

22500 355

34900 472

1.32

1.32

1.18

1.28

1.29

9660

9662

17203

9575

10966

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Figure A-5a. Wavenumber Variation of TBL Pressure SpectralDensity (F = 50 Hz)

164

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165

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Flight Condition/]I'roquency/

Normalized Froquen_

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Figure A-Sc. Wevenumber Variation of TBL Pressure SpectralDensity (F = 100 Hz)

166

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..................................................T......................_....................._................_........................fi:_..........>_:-:_:A_._............-;_H-;T-! i :: i i i !_ / _' i . 0.12711E+00

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167

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B Partial Coherence Techniques

B.1 Mathematical Formulation

The noise at a specified cabin location can be modelled as a multiple-input, single

output (MISO) system. The signals measured by microphones and accelerometers

on the aircraft surfaces represent the inputs and the signal measured by an interior

microphone represents the output. The physical measurements required to analyze

a MISO system are the power spectra of all the inputs and the output, plus all the

cross-spectra between them.

The basic frequency domain equation for a MISO system with q inputs is

q

Y(f) - y_ I-I_u(f) X_(f) + N(f) (1)i=1

where

X_(f) = Fourier transform of i th input zi(t)

Y(f) -- Fourier transform of output y(t)

Hi_(f) = Transfer function between the i th input zl and the output y

N(f) = Background Noise in the system

The auto- and cross-spectra of these time series data are calculated as

0_(/) =

O_,,.,,(/)=C_,(/)=

G.,_(1)= o,_(/) =

T._.oo_Elim2 [[y(f)[2]

lim 2T_® _e [x. (.[)xj(l)]

hm 2__.®vetx.(f)Y(/)]

where

= Expectation operator

Gii = Auto-spectrum of input zi(t)

Gvv = Auto-spectrum of output v(t)

Gij = Cross-spectrum between inputs zi(t) and xj(t)

(2)

(3)

(4)

One very important concept in multiple input and output systems is the coherence

function. It is a measure of the degree of relatedness of two signals; it is, however, not

a cause and effect relationship.

171

Ordinary" coherence function:

ICijl2G, Gj_ (5)

Partial coherence function: The partial coherence function is a measure of the

coherence between the i th selected input with the effects of all v inputs preceding it in

order removed and the jth selected output with the effects of the same r inputs removed.

la,j7q.,!- Gii.,,G_j.,, (6)

Conditioned spectrum between the i th and jth inputs:

Gij._! = Gij.(_-x)!- G'J('-I)!GI_'(_-D!G,_.(,-1)r

(7)

i th conditioned output spectrum: The conditioned spectrum is the auto-spectrum

of the i th output with the effects of the selected r input and all inputs that precede it

in order removed. The conditioned spectrum depends on the ordering of the inputs.

2 (8)Gii._! = Gii'(_-D!- G,_.(,-1)!

= Gii.(.-i), (1 -%2i.(._i),) (9)

The equations above show the relationship between the present stage of the conditioned

spectrum and the past stage of the conditioned spectrum. By induction, the following

relation between the present conditioned spectrum and its entire past spectrum can be

derived, in the form of partial coherence.

G,.., = Gi, fi (1 --"[i}r_k+l,.(r_k,O (10)

k=l

Multiple coherence function: The multiple coherence function is a measure of how

nmch energy at the output is accounted for by the inputs.

2 _ 27_.,, = 1 l:I (1- 7iu.(i-,),) (11)i=1

Eq. (7) is an iterative algorithm to compute the conditioned spectrum. It can also

be written in terms of the partial coherence function as follows.

2 7i2J(_- 1)! 1 Grj.(__I)!Gi_.(,_U ! 2"[ij.r! = (1 2 ---- z-- _'ri.(r-1)!) (1 -- "/rj.(r-1)!) G_r.(,.-i),Gij.(,.-i)!

(12)

172

Rewriting eq. (10) in dB form yields

GdS d8 dsii.r! = Gii + mii.r!

where

(13)

GidlB = 101oglo ( (Jii / (14)

AaBii.,-, = 10_l°g:to( 1 -- 7(2._k+,)i.(,._h),) (15)k=l

G,,! = 20 x 10 -6 (16)

dBIf r --_ cx_, the auto-spectrum Gii.,! will approach the background noise level.

As an example, consider the case of two inputs, when the output is denoted as y.

Go.o! = G0 (17)

Giwo! = Giu (18)

Gvv.o! = G_ (19)

For r = 0:

Forr= 1:

For r = 2:

Gij.t!

Giwl_

Gyl/.I!

= Gij.1 (20)

GljGil (21)= Gi_ GII

= Gi_.l (22)

GI_Gil (23)= Giu Gn

= G_.I (24)

GluG_I (25)= G_ Gn

IG,_,[2(26)

= G_ Gn

= Gy_ (1 - 712,) (27)

: Giy.12

: Giy.l!Gly.l_Gi1,1!

G22.1_

(28)

(29)

173

Gtty.2_ 61/i/.12

= G_.I (1-7_y.1)

= G_ (1 - 7_) (1 - 7_y.1)

2= C_(1-%u )- G,,(i-7_,)

= GI,2, (1 - 71=_,(1 - %2, (1 - I

G2vGll 2G21

GlvO2,

(30)

(31)

(32)

(33)

(34)

Note that G_v.2! is a function of combinations of input and output auto-spectra,

cross-spectra and their respective coherence functions.

For general input r > 2, the conditioned output spectrum is given as

G_.,! = G_ fl (1- 7_(,--k+l).(,-k)!)k=l

It is obvious that Gvu.,! becomes rather complicated when r is large.

(35)

B.2 Flight Test Applications

An in-house computer program was used to calculate the ordinary and partial co-

herence values and the conditioned spectra using the iterative technique outlined in

Bendat and Piersol's "Random Data - Analysis and Measurement Procedures," Chap-

ter 7, 2nd Edition (Wiley Interscience, 1986). All the processing was performed with

a resulting bandwidth of 1.5625 Hz, and yielded adequate spectral resolution without

excessive computer storage requirements.

As an example of the results of the conditioning process, Figure B-1 shows an

unconditioned output spectrum, and subsequent calculated spectra conditioned with

selected inputs (the microphone locations are shown in Figure 8-7 in the main report

text). The reduction in BPF level with each conditioning step can be seen in the figure.

174

fi Ji ! I i I_! /F i i [ i L I

-_- = , -r i i --- ii (a) Unconditioned Spectrum

t : i i ----

$

,/; _ e_'j'l

'i_BPF_'_

0

! + I i t ! I

I00 200 300 400 r._ 600 ?O(l _ gO0 tO00

Fr equenc_ (Hz)

II

II

0

.... i,+i .... !.... ! +i+c+i +i.... ' ....II +i....+..+_u_..t.............L.........+...........+........+............+.... ++++ +°+,,=++ + +i (b) Conditioned with_JV_'_'l/ViT;-i--i--- Inpu,Microphone4 T.......

..........++......i']'- ]"".........._'_ ._ .,l+......... +:.......... +i..........+............."+..........

...........i.......i : .......+...... '___.'__._ .... l i..... i...... i ........

, + + i i+,,',+!BPF ! i ', :

(reducedby lSdB <ompi_re¢lio'(t_)) ...]

I00 200 _ tO0 _ _ 700 800 900 I0O0

F'requenc _ IHzl

ft

o

- $

o

m

" " I " " " i + " T .... P .... ! " i " ! : " + " " " r ....

, i l i ] i i[ I t +

.....il .....A+.........+....._.....+......+..................i...............+.............+.........,_,i/L II7 i (c) Sp'ectrtm, conditioned with('_]t_.l]; ............i---i---i .... Input Microphones 4 and 13 ....

......... r.......+i+ -,i,,++ ............... -7....... f..........+...............+ ........i i _ l 7

.........!...........!............i......... +__ ..... ! ....!+ . ]l ' + _ i iBPF ]

-+ , ........... t ........... i ........ ;- ......... ,; _. _ . l|vv

redtlced by 21.5 dB compared to (ti)) ] t4 ! + I I ' i i

--_ ................ i ....... _ .... i ....._m 2m _o +oo soo rm _oo mo _oo mm

Frequenc_ [HI)

See Figure 8-7 for microphone locations

FIGURE B-1.

Example of the Partial Coherence ConditioningProcess for Output Microphone 5,

8x8 Test Point 19550A03

175

,IJASA Report Documentation Page

,. Report No.

NASA CR-181897

2. Government Accession No.

4. Title and Subtitle

UHB Demonstrator Interior Noise Control FlightTests and Analysis

7. Author(s)

M. A. Simpson, P. M. Druez, A. J. Kimbrough,M. P. Brock, P. L. Burg_, G. P. Mathur, M. R. Cannon,and B. N. Tran

9. Pedorming Organization Name and Addre_

Douglas Aircraft CompanyMcDonnell Douglas CorporationLong Beach, CA 90846

12, Sponsoring Agency Name and Address

National Aeronautics and Space AdministrationLangley Research CenterHampton, VA 23665-5225

3. Recipient's Catalog No.

5. Report Date

October 1989

6. Performing Org|ntz|tlon Code

8. Performing Organization Report No.

10. Work Unit No.

535-03-11-03

11. Contract or Grant No.

NAS1-18037

13. Ty_ of Report a_ P_iod Cover_

Contractor Report

14. Sponsoring _gency Code

15. Supplementary Notes

Langley Technical Monitor:Final Report - Task 4

Kevin P. Shepherd

16. A_tract

This report describes the measurement and analysis of MD-UHB Demonstrator noise andvibration flight test data related to passenger cabin noise. The objectives of thes(

analyses were to investigate the interior noise characteristics of advanced turboprol

aircraft with aft-mounted engines, and to study the effectiveness of selected noise

control treatments in reducing passenger cabin noise.

The UHB Demonstrator is an MD-80 test aircraft with the left JT8D engine replaced

with a prototype UHB engine. For these tests, the UHB engine was a General Electric

Unducted Fan, with either 8x8 or 10x8 counter-rotating propeller configurations.

Interior noise level characteristics were studied for several altitudes and speeds,

with emphasis on high altitude (35,000 ft), high speed (0.75 Mach) cruise conditionsThe efffectiveness of several noise control treatments was evaluated based on cabin

noise measurements. The important airborne and structureborne transmission paths

were identified for both tonal and broadband sources using the results of a sound

intensity survey, exteriorand interior noise and vibration data, and partial coherer

analysis techniques. Estimates of the turbulent boundary layer pressure wavenumber-

frequency spectrum were made, based on measured fuselage noise levels.

17. Key Words (Sugg_t_ by Author(s))

Acoustics

Advanced Turboprop Aircraft

Flight TestsInterior Noise Control

UHB Demonstrator

18. Distribution S_tement

Unclassified-Unlimited

Subject Category -71

19. Security Classif. (of this report)

Unclassified

20. Security Classif. (of this page)

Unclassified

21. No. of pages

190

22. Price

ce

NASA FORM 1628 OCT 86

UHB Demonstrator Interior Noise Control Flight Tests and ...€¦ · intensity survey, and Robert...

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