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I would like to acknowledge the financial support of the University of Florida's. College of Engineering and Department of Mechanical and Aerospace ...

DESIGN, FABRICATION, AND CHARACTERIZATION OF AN ANECHOIC WIND TUNNEL FACILITY

By JOSE MATHEW

A DISSERTATION PRESENTED TO THE GRADUATE SCHOOL OF THE UNIVERSITY OF FLORIDA IN PARTIAL FULFILLMENT OF THE REQUIREMENTS FOR THE DEGREE OF DOCTOR OF PHILOSOPHY UNIVERSITY OF FLORIDA 2006

ACKNOWLEDGMENTS I would like to acknowledge the financial support of the University of Florida’s College of Engineering and Department of Mechanical and Aerospace Engineering and the financial support of a NASA Langley Research Center Grant NAG1-03044, monitored by Dr. Mehdi Khorrami. I would like to thank my advisor, Dr. Louis N. Cattafesta III, for his continual guidance and motivation in making this work possible. I also would like to express my heartfelt gratitude to Dr. Mark Sheplak, Dr. Bruce Carroll, Dr. Toshi Nishida and Dr. Siddharth Thakur for their ideas and encouragement. I am also deeply indebted to Mr. Chris Bahr for his help and support. I would also like to thank numerous individuals for their invaluable contributions to this project, including Cesar Moreno, Michael Sytsma, Nik Zawodny, Ryan Holman, Todd Schultz, Ed Duell, Dragos Vieru, Raj Vaidyanathan, David Weiner, Jared Lee, Ron Brown, Wayne Willis, and Grant Pettit. My parents and sisters deserve special credit for giving me moral support and motivating me through the course of my research work. Finally, I thank God for giving me an opportunity to enjoy life as a successful doctoral student.

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TABLE OF CONTENTS Page ACKNOWLEDGMENTS .................................................................................................. ii LIST OF TABLES............................................................................................................. vi LIST OF FIGURES .......................................................................................................... vii ABSTRACT.......................................................................................................................xv CHAPTER 1

INTRODUCTION .......................................................................................................1 Background .................................................................................................................... 2 Airframe Noise ............................................................................................................... 3 Existing Anechoic Wind Tunnels................................................................................... 8 Motivation .................................................................................................................... 11 Thesis Objectives ......................................................................................................... 12 Technical Approach ..................................................................................................... 14 Thesis Organization...................................................................................................... 14

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ANECHOIC CHAMBER ..........................................................................................16 Facility Description ...................................................................................................... 16 Free Field Characterization .......................................................................................... 20 Jet Noise Characterization............................................................................................ 23

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DESIGN OF THE ANECHOIC WIND TUNNEL ...................................................33 Design Criteria ............................................................................................................. 33 Overall Layout.............................................................................................................. 35 Settling Duct/Honeycombs/Screens ............................................................................. 39 Contraction ................................................................................................................... 42 Test Section .................................................................................................................. 47 Diffuser......................................................................................................................... 48 Corner/Turning Vanes.................................................................................................. 56 Vibration Isolator ......................................................................................................... 60 Transition...................................................................................................................... 61 Fan ............................................................................................................................. 63 iii

Acoustic Treatment ...................................................................................................... 66 4

FABRICATION OF THE WIND TUNNEL COMPONENTS.................................70 Inlet ............................................................................................................................. 70 Diffuser......................................................................................................................... 71 Corner/Turning Vanes.................................................................................................. 74 Vibration Isolator ......................................................................................................... 76 Transition...................................................................................................................... 77 Fan ............................................................................................................................. 78 Acoustic Treatment ...................................................................................................... 80

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EXPERIMENTAL METHODS.................................................................................83 Chamber Deflection and Wall Loading........................................................................ 83 Tunnel Circuit Static Pressure...................................................................................... 85 Flow Uniformity........................................................................................................... 87 Shear Layer Growth ..................................................................................................... 88 Freestream Turbulence Measurements......................................................................... 91 Background Noise ........................................................................................................ 93 Fan Noise Attenuation.................................................................................................. 95 Background Noise Source Identification ..................................................................... 98 Vibration Measurements ............................................................................................ 100 Acoustic Liner Absorption Coefficient Measurement Setup ..................................... 102

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FACILITY CHARACTERIZATION ......................................................................104 Chamber Deflection and Wall Loading...................................................................... 104 Inlet Wall Pressure ..................................................................................................... 106 Diffuser Wall Pressure ............................................................................................... 110 Flow Uniformity......................................................................................................... 112 Shear Layer Behavior................................................................................................. 115 Freestream Turbulence ............................................................................................... 120 Background Noise Measurements.............................................................................. 126 Fan Noise Decay ........................................................................................................ 132 Background Noise Source Identification ................................................................... 135 Vibration Measurements ............................................................................................ 143 Acoustic Liner Absorption Coefficient Estimation.................................................... 150

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CONCLUSIONS AND FUTURE WORK ..............................................................152

REFERENCES ................................................................................................................156 APPENDIX

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A

SCHEMATICS OF THE WIND TUNNEL ............................................................162

B

DERIVATION OF THE INLET SHAPE POLYNOMIAL ....................................164

C

INLET OPTIMIZATION STUDY ..........................................................................169

D

DIFFUSER OPTIMIZATION STUDY...................................................................174

E

FAN LOSS CALCULATION .................................................................................179

F

EFFECT OF LEAKAGE ON WALL PRESSURE .................................................185

G

RESULTS OF FREE FIELD CHARACTERIZATION..........................................192

BIOGRAPHICAL SKETCH ...........................................................................................201

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LIST OF TABLES page

Table

1-1. Details of the existing anechoic wind tunnels. ..........................................................11 2-1. SPL deviation errors for northeast directions, before and after treatment.................23 3-1. Summary of the wind tunnel design. .........................................................................38 3-2. Test section details.....................................................................................................47 3-3. Turning vane coordinates. .........................................................................................60 3-4. Results of the wind tunnel circuit loss calculation. ...................................................63 6-1. Axial location of the inlet pressure taps. ...................................................................86 5-1. Location of the diffuser microphones .......................................................................98 6- 2. Spectral error estimates. .........................................................................................125 6-3. Free Stream Turbulence Intensity............................................................................126 6-4. Error estimates for the background noise spectra....................................................128 6-5. Nomenclature for input and output microphones. ...................................................137 C-1. Results of Inlet optimization study. ........................................................................172 D-1. Final dimensions of the tunnel obtained from the optimization study. ..................178

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LIST OF FIGURES page

Figure

1-1. Aircraft noise sources. .................................................................................................3 1-2. Propulsive noise reduction through the ages. ..............................................................4 1-3. Relative magnitudes of the various aircraft noise components during landing...........5 1-4. Various airframe noise sources....................................................................................6 2-1. Schematic of the original UF anechoic chamber.......................................................18 2-2. Schematic of the wall wedges. ..................................................................................19 2-3. Cross sectional view of the chamber wall panel........................................................19 2-4. Measurement array paths in the anechoic chamber...................................................21 2-5. Deviation of pressure measurements from free field from chamber center towards bell-mouth. ...............................................................................................................22 2-6. Deviation of pressure measurements from free field from chamber center towards Northeast corner of room with double door. ............................................................22 2-7. Side view schematic of the jet noise measurements..................................................25 2- 8. Top view schematic of the jet noise measurements. ................................................26 2-9. Schematic of the jet nozzle........................................................................................26 2-10. Cold jet noise data measured at 90o to the jet axis at 83.5 jet diameters for various jet Mach numbers. Exhaust fan is off. Compressor on. ..........................................28 2-11. Cold jet noise data measured at 90o and 140o to the jet axis at 83.5 and 114.5 jet diameters, respectively, at M=0.9. Exhaust fan is off. F and G are the large- and fine-scale similarity third-octave band spectra ........................................................28 2-12. Cold jet noise data measured at 90o to the jet axis at 83.5 jet diameters for various jet Mach numbers. Exhaust fan is operating at max speed (~6000 CFM). .............30

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2-13. Comparison between blowdown (compressor off) and continuous (compressor on) operating conditions for approximately identical flow conditions (M=0.7±0.01) (measured at 90o to the jet axis). ..............................................................................30 3-1. Wind tunnel design flow chart...................................................................................35 3-2. Plan view of the wind tunnel. ....................................................................................37 3-3. Schematic of the honeycomb section. .......................................................................40 3-4. Schematic of the screen design..................................................................................42 3-5. C p distribution along corner for a contraction..........................................................43 3-6. Schematic of the contraction shape polynomial. .......................................................44 3-7. Contours of x velocity along the half mid-plane for the contraction.........................46 3-8. Schematic of the collector. ........................................................................................49 3-9. Schematic of the 2D diffuser. ....................................................................................49 3-10. 2-D Diffuser Design Curves. ...................................................................................51 3-11. Comparison of local pressure coefficient with Stratford’s separation pressure coefficient for diffuser 1...........................................................................................54 3-12. Comparison of local pressure coefficient with Stratford’s separation pressure coefficient for diffuser 2...........................................................................................54 3-13. Centre plane x velocity profile along diffuser 1. .....................................................55 3-14. Schematic of the Turning Vanes. ............................................................................58 3-15. Results from turning vane simulation for a test section speed of 76 m / s ..............58 3- 16. Results from turning vane simulation for a test section speed of 18 m / s .............59 3-17. Schematic of the rectangular to round transition section ( H e = 2.22 m , W = 1.2 m , D = 1.95 m )..............................................................................................................62 3-18. Results from transition flow simulation for a test section speed of 76 m / s ..........62 3-19. Fan Load curve. .......................................................................................................65 3-20. Estimated pressure drop along the wind tunnel circuit for a test section velocity of 76 m / s .....................................................................................................................65

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3-21. Details of the wind tunnel duct walls. .....................................................................69 4-1. Photograph of the inlet contraction. ..........................................................................71 4-2. Photograph of diffuser 1. ...........................................................................................72 4-3. Diffuser 1 internal skeletal view................................................................................72 4-4. Cross sectional view of the ‘I’ beam. ........................................................................73 4-5. Structural reinforcement using polyurethane foam. ..................................................73 4-6. Structural reinforcement using semi cylindrical hollow fiberglass sheets. ...............74 4-7. Photograph of the turning vane rack..........................................................................75 4-8. Side view of the cross plate. ......................................................................................75 4-9. Photograph of vane mold...........................................................................................76 4-10. Photograph of the vibration isolator section............................................................77 4-11. Photograph of the transition piece. ..........................................................................78 4-12. Front view of the fan. ..............................................................................................79 4-13. Back view of the fan................................................................................................79 4-14. View of the fan base. ...............................................................................................80 4-15. Chamber traverse acoustic treatment.......................................................................81 4-16. Garage door acoustic treatment. ..............................................................................82 4-17. Photograph of the flow silencer...............................................................................82 5-2. Schematic of the chamber wall loading measurement setup.....................................84 5-3. Schematic of the inlet static pressure taps. ................................................................86 5-4. Photograph of the inlet static pressure taps. ..............................................................87 5-5. Schematic of the flow uniformity measurement setup. .............................................88 5-6. Schematic of the shear layer measurement setup. .....................................................90 5-7. Photograph of the shear layer measurement setup. ...................................................90 5-8. Hotwire measurement block diagram........................................................................93

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5-9. Photograph of the background noise measurement setup. ........................................94 5-10. Coherent power measurement. ................................................................................95 5-11. Schematic of the fan noise measurement microphone holder. ................................97 5-12. Setup for measurement of fan noise decay..............................................................97 5-14. Photograph of the fan vibration measurement test arrangement. ..........................101 5-15. Photograph of the vibration isolator vibration test arrangement. ..........................102 5-16. Schematic of the Impedance tube setup. ...............................................................103 6-1. Wall loading as a function of test section speed......................................................105 6-2. Variation of effective velocity with the test section velocity. .................................105 6-3. Wall deflection vs test section velocity. ..................................................................106 6-4. Contraction C p distributions versus length for the (a) sidewall, (b) base, and (c) corner for U TS = 17 m / s . .......................................................................................107 6-5. Contraction C p distributions versus length for the (a) sidewall, (b) base, and (c) corner for U TS = 30 m / s . .......................................................................................108 6-6. Contraction C p distributions versus length for the (a) sidewall, (b) base, and (c) corner for U TS = 42 m / s ........................................................................................108 6-7. Comparison of the pressure drop across the inlet and flow conditioner section for a test section speed of a) 18 m / s b) 37 m / s ..........................................................109 6-8. Pressure recovery across the diffuser. .....................................................................111 6-9. Photograph showing the waviness of the inner surface of diffuser 2......................111 6-10. Comparison of the pressure recovery across the diffuser duct work section for a test section speed of a) 18 m / s b) 37 m / s ...........................................................112 6-11. Normalized stagnation pressure contours (max =1 w/ 0.1 interval) at the test section entrance. H e and We are the height and the width at the diffuser 1 entrance for a test section speed of 17 m / s . ........................................................................113 6-12. Normalized stagnation pressure contours (max =1 w/ 0.1 interval) at the diffuser entrance for a test section speed of 17 m / s ...........................................................114

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6-13. Test section centerline velocity profile development along the test section length for a test section speed of 17 m / s . ........................................................................114 6-14. Normalized velocity profile in the yz plane in the y direction for U TS = 18 m / s ( θ1 = 7.8 mm ). ........................................................................................................116 6-15. Normalized velocity profile in the yz plane in the y direction for U TS = 30 m / s ( θ1 = 6.9 mm ).........................................................................................................116 6-16. Normalized velocity profile in the zy plane in the y direction for U TS = 37 m / s ( θ1 = 7.5 mm ). ..................................................................117 6-17. Normalized velocity profile in the yz plane in the z direction for U TS = 18 m / s ( θ1 = 7.1 mm ). ........................................................................................................117 6-18. Normalized velocity profile in the yz plane in the z direction for U TS = 30 m / s ( θ1 = 7.2 mm ).........................................................................................................118 6-19. Normalized velocity profile in the yz plane in the z direction for U TS = 37 m / s ( θ1 = 7.2 mm ).........................................................................................................118 6-20. Variation of y momentum thickness with test section length for a) U TS = 18 m / s b) U TS = 30 m / s c) U TS = 37 m / s .............................................................................119 6-21. Variation of z momentum thickness with test section length for a) U TS = 18 m / s b) U TS = 30 m / s c) U TS = 37 m / s .............................................................................119 6-22. Variation of the potential core velocity along the test section length for a) U TS = 18 m / s b) U TS = 30 m / s c) U TS = 37 m / s .................................................120 6-23. Locations of the hotwire measurement..................................................................121 6-24. Calibration curve showing the plot of mean velocity vs. mean voltage................122 6-25. Calibration curve corrected for flow temperature. ................................................122 6-26. Cubic fit to the calibration curve. ..........................................................................123 6-27. Turbulence spectra at location A. ..........................................................................123 6-28. Turbulence spectra at location B. ..........................................................................124 6-29. Turbulence spectra at location C. ..........................................................................124

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6-30. Narrow-band Spectra.............................................................................................127 6-31. 1/3rd Octave Band Spectra. ....................................................................................127 6-32. OASPL vs test section velocity .............................................................................128 6-33. Comparison of UF and Notre Dame tunnel background noise. ............................129 6-34. Narrow band inflow spectra. .................................................................................131 6-35. The influence of inflow microphone on the outflow spectra.................................131 6-36. Facility comparison of A-weighted in flow noise levels.......................................132 6-37. Total power measured by the diffuser 2 microphones for a) U TS = 18 m / s b) U TS = 30 m / s c) U TS = 42 m / s . ...........................................................................133 6-38. Total power measured by the diffuser1 microphones for a) U TS = 18 m / s b) U TS = 30 m / s c) U TS = 42 m / s . ...........................................................................133 6-39. Total coherent power measured by the diffuser2 microphones for a) U TS = 18 m / s b) U TS = 30 m / s c) U TS = 42 m / s . .......................................................................134 6-40. Total coherent power measured by the diffuser1 microphones for a) U TS = 18 m / s b) U TS = 30 m / s c) U TS = 42 m / s . .......................................................................134 6-41. Schematic of the MISO model. .............................................................................136 6-42. Autospectra of the input and output microphones for U TS = 18 m / s . ..................138 6-43. Autospectra of the input and output microphones for U TS = 30 m / s . ..................138 6-44. Autospectra of the input and output microphones for U TS = 42 m / s ...................139 6-45. Ordinary coherence between the input microphones and output microphone for U TS = 18 m / s . ........................................................................................................139 6-46. Ordinary coherence between the input microphones and output microphone for U TS = 30 m / s .........................................................................................................140 6-47. Ordinary coherence between the input microphones and output microphone for U TS = 42 m / s .........................................................................................................140 6-48. Comparison of the MISO model to the measured spectra for U TS = 18 m / s . ......141

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6-49. Comparison of the MISO model to the measured spectra for U TS = 30 m / s . ......141 6-50. Comparison of the MISO model to the measured spectra for U TS = 42 m / s .......142 6-51. Total power for model and measured output for U TS = 18 m / s . ..........................142 6-52. Total power for model and measured output for U TS = 30 m / s . ..........................143 6- 53. Total power for model and measured output for U TS = 42 m / s ..........................143 6-54. Autospectra of the accelerometers attached to a) Fan slab b) Retainer wall for U TS = 18 m / s . ........................................................................................................144 6-55. Autospectra of the accelerometers attached to a) Fan slab b) Retainer wall for U TS = 30 m / s .........................................................................................................145 6-56. Autospectra of the accelerometers attached to a) Fan slab b) Retainer wall for U TS = 42 m / s .........................................................................................................145 6-57. Transmission loss across the fan base for the x axis accelerometer for a) U TS = 18 m / s b) U TS = 30 m / s c) U TS = 42 m / s . ...............................................146 6-58. Transmission loss across the fan base and the building floor for the x axis accelerometer for a) U TS = 18 m / s b) U TS = 30 m / s c) U TS = 42 m / s . ..............146 6-59. Autospectra of the accelerometers attached to the vibration isolator for U TS = 18 m / s . ........................................................................................................148 6-60. Autospectra of the accelerometers attached to the vibration isolator for U TS = 30 m / s .........................................................................................................148 6-61. Autospectra of the accelerometers attached to the vibration isolator for U TS = 42 m / s .........................................................................................................149 6-62. Transmission loss across the vibration isolator for the x axis accelerometer for a) U TS = 18 m / s b) U TS = 30 m / s c) U TS = 42 m / s . ...............................................149 6-63. Time response of a turning vane due to an impulsive impact. ..............................150 6-64. Normal incidence absorption coefficient for the acoustic liner.............................150 A-1. Plan view of the wind tunnel. .................................................................................162 A-2. Cross-section view of the wind tunnel....................................................................163

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A-3. North elevation of the wind tunnel. ........................................................................163 B-1. Schematic of the inlet shape polynomial. ...............................................................164 B-2. Plot of contraction shape polynomial in the x-y plane. ..........................................168 B-3. Plot of contraction shape polynomial in the x-z plane............................................168 C-1. Velocity vector at the inlet exit plane. ....................................................................169 D-1. Wind tunnel flow path. ...........................................................................................175 D-2. Location of diffuser 1 and 2 designs on the Kline’s flat diffuser curves................178 F-1. Schematic of the chamber. ......................................................................................185 F-2. Equivalent electric circuit representation of the chamber flow...............................187 F-3. Variation of leakage ratio with the leakage flow resistance....................................190 F-4. Variation of the wall pressure differential with leakage area ratio for various leakage resistance ratios. .....................................................................................................191 F-5. Variation of the wall force with leakage area ratio for various leakage resistance ratios. ......................................................................................................................191 G-6. Deviation of pressure measurements from free field from chamber center in the Northeast direction. ................................................................................................193 G- 7. Deviation of pressure measurements from free field from chamber center in the West direction. .......................................................................................................194 G-8. Deviation of pressure measurements from free field from chamber center in the Northwest direction. ...............................................................................................195 G-9. Deviation of pressure measurements from free field from chamber center in the North direction. ......................................................................................................196 G-10. Deviation of pressure measurements from free field from chamber center in the Southwest direction. ...............................................................................................197 G-11. Deviation of pressure measurements from free field from chamber center in the South direction. ......................................................................................................198 G-12. Deviation of pressure measurements from free field from chamber center in the Southeast direction. ................................................................................................199 G-13. Deviation of pressure measurements from free field from chamber center in the East direction..........................................................................................................200 xiv

BSTRACT

Abstract of Dissertation Presented to the Graduate School of the University of Florida in Partial Fulfillment of the Requirements for the Degree of Doctor of Philosophy DESIGN, FABRICATION, AND CHARACTERIZATION OF AN ANECHOIC WIND TUNNEL FACILITY By Jose Mathew May 2006 Chair: Louis Cattafesta Major Department: Mechanical and Aerospace Engineering The design, fabrication, and characterization of an anechoic wind tunnel facility at the University of Florida are presented. The objective of this research is to develop and rigorously characterize an anechoic wind tunnel suitable for detailed aerodynamic and aeroacoustic research. A complete tunnel design methodology is developed to optimize the design of the individual components of the wind tunnel circuit, and modern analysis tools, such as computational fluid dynamics and structural finite element analyses, are used to validate the design. The wind tunnel design is an “L-shaped”open circuit with an open jet test section driven by a 300 HP centrifugal fan. Airflow enters the wind tunnel through a settling duct with a honeycomb section and a set of four screens. An optimized, minimum length (3.05 m) 8:1 contraction accelerates the flow into a rectangular test section that measures 0.74 m by 1.12 m by 1.83 m. Mach number similarity dictates the maximum velocity xv

attainable in the test section to be 76 m/s; thus the maximum Reynolds number based on chord (chord=2/3 span) attainable is in the 3-4 million range. The flow leaving the test section enters an acoustically treated and 2D diffuser that simultaneously provides static pressure recovery and attenuates fan noise. The flow then turns a 90° corner with turning vanes and enters a second diffuser. The flow leaving the second diffuser enters the fan through a transition section. The wind tunnel was characterized rigorously at speeds up to 43 m/s to ensure the quality of the future aerodynamic and aeroacoustic measurements. The overall SPL from 100 Hz – 20 kHz ranges from 54.8 dB at 18 m / s to 75.7 dB at 43 m / s . The freestream turbulence level has a value of 0.035 %, and the flow non uniformity in the test section was found to be < 0.7 % for a test section speed of 17 m / s . The outcome of this work is an anechoic wind tunnel with excellent flow quality, low background noise, and the largest Reynolds number capability among universityscale anechoic facilities in the US.

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CHAPTER 1 INTRODUCTION The goal of this research is to design, fabricate, and characterize an anechoic wind tunnel facility at the University of Florida. An existing anechoic facility at the University of Florida (Jansson et al. 2002) has been upgraded to an anechoic wind tunnel. The purpose of this endeavor is to permit high quality fluid dynamic and aeroacoustic experiments on airframe noise, with an initial focus on trailing edge noise. A research flow facility with low turbulence levels, good flow uniformity, and low background noise levels that can achieve high chord Reynolds numbers in the test section is essential in this regard. Strict regulations imposed by the Federal Aviation Authority (FAA, 2004) on the noise from commercial aircraft have increased the emphasis on airframe noise, which is a significant portion of the aircraft noise during approach and landing. A reduction in aircraft noise will require attenuation of airframe noise produced by specific aircraft components, such as airfoil trailing edges, landing gear, airfoil flaps and slats, and wing tips. A fundamental understanding of the noise generation mechanisms will provide the ability to model and predict the emitted noise and may enable researchers to devise effective schemes to ultimately reduce airframe noise. However, appropriate experiments conducted in an anechoic wind tunnel are required to achieve significant advances in this regard. This chapter presents an overview of airframe noise and its various components, a survey of other existing anechoic flow facilities, the motivation for this research, the technical objectives and approach, and also provides an outline of this thesis. 1

2 Background Commercial air traffic has been growing at such a fast pace that aircraft noise has increasingly become an annoyance to the communities in close proximity to airports, and the importance of aircraft noise reduction is now being realized by the international community (Willshire 2001). The detrimental effects of aircraft noise include sleep deprivation, irritability, reduced land resale value, and the delay in the growth of civil aviation. Realizing these drawbacks, NASA has proposed a plan as a part of the Quiet Aircraft Technology (QAT) program “to reduce the perceived noise levels of future aircraft by 10 (decibels) dB from today’s subsonic aircraft and by 20 dB within twenty five years” (Goldin 1997). Figure 1-1 shows the various sources of aircraft noise. Aircraft noise consists mainly of airframe noise and power plant noise. Power plant noise includes jet noise, turbomachinery noise, and combustion noise, while airframe noise consists of noise due to flaps, slats, landing gear, wing and tail, etc. Considerable research over the past three decades has focused on the reduction of aircraft jet noise.

Figure 1-2 (www.aia-

aerospace.org, 2004) shows the progress achieved in the development of quieter civilian aircraft over the past 50 years. The turbofan engines of the present day are at least 20 dB quieter than the turbojet engines of the early sixties. The use of ultra-high bypass ratio turbofan engines and the successful implementation of liner technology has helped mitigate jet noise to such an extent that airframe noise or non-propulsive noise has now become a significant source of aircraft noise (Crighton 1995), especially during approach and landing. The relative magnitudes of the various aircraft noise components during approach are shown in Figure 1-3.

3

AIRCRAFT NOISE

AIRFRAME NOISE

Flap

Slat

POWER PLANT NOISE

Landing Gear

Wing/ Tail

Jet Noise

Turbomachinery Noise

Combustion Noise

Figure 1-1. Aircraft noise sources. During takeoff, the aircraft engine is operating at maximum thrust, and therefore jet noise and fan exhaust noise dominates over other noise sources.

However during

approach the aircraft engine is flying at low power and all the high lift devices and the landing gear are fully extended, resulting in a greater contribution of airframe noise to the total aircraft noise spectrum. Note that during approach, airframe noise is comparable to the fan inlet noise, making them the primary noise sources during approach. In order to design quieter airplanes, the physics behind the various airframe noise generation mechanisms must be thoroughly understood. Airframe Noise Airframe noise is defined as the total aircraft noise minus the noise from the engine and noise from engine-airframe interference (Lockard and Lilley 2004). Various sources of airframe noise are annotated in Figure 1-4 (Golub et al. 2004).

4

Lateral N oise Level C orrected for A ircraft Thrust

Turbojets First Generation Turbofans Second Generation Turbofans

20 dB

1960

1965

1970

1975

1980

1985

1990

1995

2000

Entry into Service Date

Figure 1-2. Propulsive noise reduction through the ages. The main sources of airframe noise are the flaps, slats and landing gear. Noise also emanates from fuselage, wing, tail, landing gear cavities, etc. A ‘clean’ full scale airframe with flaps, slats, and landing gear retracted generates mainly broadband noise with the broadband peak located in the vicinity of several hundred Hz (Smith 1989). An aircraft during approach has its flaps, slats and landing gear extended, increasing the overall airframe noise levels by approximately 10 dB. The landing gear noise is omnidirectional and has spectral characteristics higher in frequency than the clean airframe (Smith 1989). Landing gear noise is caused in part by vortex shedding over bluff bodies like wheels, axles, struts, shafts, etc. (Crighton 1995). At the typical shedding frequency, the sound radiated has a U 6 dependence on velocity, where

5 U is a typical velocity in the flow field. High-lift devices like wing flaps and slats

modify the spectral content of the clean airframe, tending to lower its characteristic frequency as a result of extending the chord of the wing and inducing larger turbulence in the wing wake (Smith 1989). Noise from high lift devices has been shown to exhibit a

U 5 dependence. P&W ADP Engine

P&W 1992 Technology Engine

Inlet

Aftfan

Combustor

Turbine

Jet

Total Airframe

Total Aircraft Noise 60

70

80

90

100

110

EPNdB

Figure 1-3. Relative magnitudes of the various aircraft noise components during landing. There is a large body of literature available on the theory of trailing edge noise from two-dimensional airfoils. Howe (1978) categorized trailing edge noise theories into three categories based on a) Lighthill’s acoustic analogy, b) linearized hydrodynamic equations, and c) ad-hoc models. All models predict a U 5 dependence of the radiated sound on the freestream velocity U . When turbulent boundary layer eddies convect past

6 the trailing edge of an airfoil ( chord, c < λ ), the acoustic scattering produces broadband

radiation to the farfield (Ffowcs Williams and Hall 1970; Crighton and Leppington 1971). However, if coherent vortex shedding is present (e.g., due to blunt trailing edges at high angles of attack), tonal or narrowband noise is also present. Khorrami et al. (2000) used highly resolved unsteady Reynolds Averaged Navier Stokes (URANS) computations of a blunt trailing edge flow to reveal strong vortex shedding and corresponding acoustic wave propagation from the trailing edge region.

Blake and

Gershfeld (1989) earlier explained how tonal noise is generated when periodic vorticity results due to an instability in the trailing edge wake. Furthermore, they identified the broadband noise spectral component, occurring due to the convection of turbulence past the trailing edge.

Vertical Tail

Slats

Flaps

Nose Landing Gear Main Landing Gear

Figure 1-4. Various airframe noise sources.

Horizontal Tail

7 Several researchers have performed experimental studies of trailing edge noise via unsteady surface pressure measurements and directional acoustic arrays. Brooks and Hodgson (1981) used measured surface pressures near the trailing edge to arrive at a correlation for the noise generated. Brooks and Marcolini (1985) used a cross-spectral technique to determine noise sources from trailing edges and used the resulting data to formulate scaling laws for trailing edge noise. More recently, Hutcheson and Brooks (2002) have compared directional array measurements to a cross spectral method that uses a pair of microphones on opposite sides of the airfoil. Macaraeg (1998) described a fundamental investigation of airframe noise using extensive flow visualization, velocity and noise measurements on a small-scale, part-span flap model. Kunze et al. (2002) have measured trailing edge noise from flat plate geometries and devised a procedure to distinguish the trailing edge noise component from the background noise. All of these theoretical and experimental results have demonstrated the importance of the following primary nondimensional parameters:

Mach number, M ∞ , chord

Reynolds number, Rec , angle of attack, c θ , t θ , (i.e., ratio of airfoil chord length, c , or trailing edge thickness, t , to local boundary layer momentum thickness, θ ). In addition, the structure of the turbulent boundary layer in the vicinity of the trailing edge (e.g., shape factor and nondimensional pressure gradient) is also important. Furthermore, practical aircraft configurations have variable wing sweep, which leads to cross flow and the development of three-dimensional boundary layers on the wing. The freestream turbulence intensity and the farfield boundary conditions are also significant parameters in experimental studies.

8 Since most previous research has focused on noise from flow over two-dimensional airfoils, much less is known about the trailing edge noise from swept wings that exist on modern commercial aircraft.

It is anticipated that trailing edge sweep will have a

profound effect on the sound generation and directivity patterns due to the threedimensional nature of the boundary layer. Our goal is to develop a anechoic wind tunnel facility that simulates a free field acoustic flight environment. The facility will enable us to conduct future benchmark experiments to eventually understand the relation among wing sweep, noise generation mechanisms, and trailing edge noise radiation patterns. Typical experiments include the measurement of surface pressure fluctuations on the trailing edge using flush mounted pressure sensors. Since wing sweep leads to the development of a three dimensional boundary layer on the airfoil, crossflow measurements of the three-dimensional boundary layer must also be made. Surveys in the wake region must be conducted to estimate the amplitude and the frequency of vortex shedding off the trailing edge. Acoustic measurements will include the measurement of amplitude and directivity of the far field radiated noise from the trailing edge. As a first step towards our goal, a high quality anechoic wind tunnel facility with low turbulence and low background noise levels must be fabricated and characterized. There are not many anechoic wind tunnels in the US where airframe measurements can be conducted (Duell et al. 2002). A survey of existing anechoic wind tunnels in the world was conducted prior to designing our wind tunnel and is summarized below. Existing Anechoic Wind Tunnels

Anechoic wind tunnels are used extensively by both the automotive and aerospace industries for scaled model testing.

While the aerospace industry has focused on

improved aeroacoustic measurements with secondary emphasis on new anechoic wind

9 tunnels, the automotive industry has focused more on developing quieter anechoic wind tunnels. The existing trend in the wind tunnel technology is to construct larger facilities with lower background noise levels (Duell et al. 2002). The larger facilities can provide higher Reynolds numbers and lower turbulence intensities in the test section, since the turbulence intensities drops with increasing the contraction ratio of the inlet. However, the cost of building these facilities is enormous, and also their power consumption is high because the power required to run the wind tunnel fan scales with the area of the test section and the third power of test section velocity (Pope & Harper 1966).

The

maintenance of large facilities is also difficult. Larger facilities also require a larger fan to operate, which in turn increases the background noise in the test section. As an example consider the Daimler Chrysler wind tunnel facility located in Detroit, which is used for automotive testing (Walter et al. 2003). The wind tunnel inlet has an entrance area of 29 m2 and the facility itself occupies an area that spans 31,000 ft 2 . The fan required to drive the facility requires a 6343 HP motor and the facility cost 37.5 million dollars to build. The characteristics of existing anechoic wind tunnels are summarized in Table 1-1. The facilities shown in the table include industrial tunnels, government tunnels and university scale tunnels. These tunnels are used for automotive component testing or aircraft component testing. The wind tunnel could be of the blower, blowdown or the suction type. There are pros and cons for each design. The wind tunnel can also be of the open circuit or the closed circuit type. For a blower tunnel the fan is located upstream of the test section, and blows high speed flow into the test section through an inlet contraction. Although the pressure in the test section is atmospheric, the flow quality is generally not optimal. The blower tunnels

10 also suffer from a low frequency pumping effect due to spillage from the collector. This low frequency phenomenon can potentially match the resonance frequency of the wind tunnel structure and even damage the structure. A blowdown tunnel is a very simple design which uses a nozzle fed by a compressed air storage tank to accelerate flow in the test section. However to achieve steady flow over a large run time, the tank size is prohibitive. A very common and efficient design is the suction tunnel, which uses a downstream fan to pull the flow through the wind tunnel circuit. Although the pressure drops to sub-atmospheric in the test section, the flow quality is generally higher than that for a blower tunnel, for equivalent amounts of flow conditioning. A closed circuit tunnel circulates the same air through the tunnel circuit, thereby turbulence can be reduced, as opposed to an open circuit tunnel, where the scales of incoming atmospheric turbulence are much larger. However, the additional duct work and space required in a closed circuit design renders the cost much higher than that for an open circuit tunnel. The quantities of most interest are the maximum Reynolds number, flow uniformity, turbulence intensity, ( u′ / U .100 ), where u′ is the root mean square value of the axial component of the turbulent fluctuations, and the background noise level. The maximum Reynolds numbers for the facilities are based on the test section hydraulic diameter. A good facility should provide high Reynolds numbers, good flow uniformity, low turbulence intensities, and low background noise levels. The background noise levels have to be very low (preferably at least 10 dB below typical levels of measurement) in an anechoic wind tunnel to make good quality acoustic measurements (Duell et al. 2002). More details of this will be given in Chapter 3.

11 Table 1-1. Details of the existing anechoic wind tunnels. Facility

Circuit Type

Drive

Test Section Type

Test Section Size (m)

Max Speed (m/s)

Max Re # (million)

Langley QFF

Open

Pressure/ Vacuum

Open

0.61 x 0.91

58

2.8

Boeing LSAF

Open

Blower

Open

2.74 x 3.66

85

17

Open

Fan

Open

2 diam

130

17

Open

Fan

Open

0.38 x 0.51

75

2.1

Fan

Closed Closed Closed

9.5 x 9.5 8x6 6x6

62 116 152

38 51 59

Open

9.5 x 9.5

85

52

ONERA CEPRA 19, France NLR, Holland DNW, Holland NUWC, Newport IVK, Stuttgart

Closed

Open

Fan

Open

1.22 diam

61

4.8

Closed

Fan

Open

5.8 x 3.87

80

24

Notre Dame

Open

Fan

Open/Closed

0.61 x 0.61

28

1.1

Audi, Germany DTF WT8, Detroit

Closed

Fan

Open

3.94 x 2.8

83

17

Closed

Fan

Nissan, Japan

Closed

Fan

Daimler Chrysler AAWT, Detroit

Closed Circuit

RTRI, Japan

Closed

Fan

Flow Uniformity

Freestream Turbulence Intensity

0.15 %

0.5%

0.3 % (velocity) 0.04% 0.3 % (velocity) 0.4 % (SP)

0.3 %

Open

5.56 x 3.34

54

14

Open

4.15 x 2.54

67

13.6

Open

4x7

53

17

0.13 %

Open

3x5

75

18

0.15 %

Open

6.9 x 4.0

71

23.4

0.25 % (SP)

Open

3 x 2.5

83

14.6

0.7 % (@ 90 m/s)

Closed

5x3

111

27

Fan

Ford, Germany

Closed

Fan

Open

20 m2

53

Virginia Tech

Closed

Fan

Closed

1.83 x 1.83

80

9.4

NASA Ames

Open

Fan

Closed

24.2 x 12.1

154

160

NASA Glenn

Closed

Fan

Closed

4.6 x 2.74

68

15

0.5 % (@ 55 m/s) 0.5 % (velocity)

0.34 %

0.16 %(@ 62.5 m/s) 0.2 % (@ 100 m/s) 0.2 % (@ 55 m/s)

Background Noise Level 33 dB (1 kHz) (@ 18 m/s) 65 dBA (Outflow) (@ 35 m/s)

80 dBA (@ 80 m/s) 71 dBA (@ 41.7 m/s) 45 dB(1 kHz, third octave) (@ 20 m/s) 60 dBA (@ 44.4 m/s) 63.7 dBA 66 dBA (@ 27.8 m/s) 62.3 dBA (@ 28 m/s) 75.6 dBA (@ 83.3 m/s)

72 dBA (@ 38.9 m/s) 75 dBA (Inflow) (@ 32 m/s) 86 dBA (Inflow) (@ 52 m/s) 65 dBA (Inflow) (@ 22 m/s)

Motivation

Wind tunnels facilities are ubiquitous, whereas few anechoic wind tunnel facilities exist in the US or rest of the world. At a university level, currently only two such facility

12 exists in the US. One of them is the subsonic, low turbulence anechoic wind tunnel at the University of Notre Dame, for which there is established data (Mueller et al. 1992). There is also a partially anechoic wind tunnel facility at Virginia Tech, which is currently being upgraded from an aerodynamic stability tunnel to an anechoic facility (Smith et al. 2005). As such, the Notre Dame facility represents the benchmark for comparison. In particular, the maximum velocity attainable in the Notre Dame wind tunnel is 28 m/s, and hence the maximum Reynolds number based on the test section hydraulic diameter is 1.1.106. The test section of the Notre Dame facility measures 0.6 m by 0.6 m (24” by 24”). The inlet contraction has a contraction ratio of 20 and it is 4.26 m long (14 ft), and is therefore rather large and expensive, but does provide freestream turbulence levels of 0.04%. For the industrial and governmental wind tunnel facilities, airframe noise measurements are often limited by scheduling and budgetary constraints. Our goal is to build a facility that offers high quality airframe noise and flow measurements at relatively low

cost

(