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Prediction of Rotorcraft Noise with A Low-Dispersion Finite Volume Scheme

A Thesis Proposal Presented to The Faculty of the Division of Graduate Studies By Gang Wang Advisor: Dr. T. C. Lieuwen. Prediction of Rotorcraft Noise with A Low-Dispersion Finite Volume Scheme. Background Approach Results Conclusions Proposed Work. OUTLINE.

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Prediction of Rotorcraft Noise with A Low-Dispersion Finite Volume Scheme

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  1. A Thesis Proposal Presented to The Faculty of the Division of Graduate Studies By Gang Wang Advisor: Dr. T. C. Lieuwen Prediction of Rotorcraft Noise with A Low-Dispersion Finite Volume Scheme

  2. Background Approach Results Conclusions Proposed Work OUTLINE

  3. Helicopter has a wide range of military and civil applications. However, the high noise level associated with it greatly restricts its further applications. BACKGROUND

  4. Three categories of rotor noise Rotational noise Broadband noise Impulsive noise High Speed Impulsive (HSI) noise Blade Vortex Interaction (BVI) noise BACKGROUND

  5. BACKGROUND High Speed Impulsive noise

  6. BACKGROUND Blade Vortex Interaction noise

  7. Many efforts have been spent on quantifying and minimizing rotorcraft noise. Three noise prediction techniques: High resolution aerodynamics in the near field and acoustic analogy for radiation in far field High resolution aerodynamics in the near field and Kirchhoff’s formula for radiation in far field Fully computational aerodynamics and acoustics BACKGROUND

  8. BACKGROUND Far Field Observer Acoustic calculation Region Blade  CFD calculation Region

  9. Much progress has been made during the past two decades in understanding and predicting rotorcraft noise characteristics with the aid of Computational Fluid Dynamics. However, dispersion and dissipation errors accompanied with conventional CFD methods alter the observed noise characteristics even a short distance away from the rotor. BACKGROUND

  10. Significant computing resources are needed to reduce these errors. This precludes the prediction methodology from use in engineering design and development. Dispersion and dissipation phenomena can be simply shown by tracking rectilinear propagation of a Gaussian sound pulse: BACKGROUND

  11. BACKGROUND Gaussian Pulse Distribution

  12. BACKGROUND T=0 T=50 Magnitude drops as wave propagates… Dissipation T=100 Dissipation Phenomenon

  13. BACKGROUND T=50 T=100 T=0 Dispersion Errors- some waves travel slower than the rest. Dispersion Phenomenon

  14. Develop an improved algorithm with low dispersion and dissipation errors. The schemes should be simple enough so that they can find immediate use in CFD codes which are widely used in industry. It should not sacrifice aerodynamic resolution for acoustic resolution, and vice versa. OBJECTIVES

  15. The integral form of Navier-Stokes equations may be written as: The flux across the cell boundary is split into two parts and : APPROACH

  16. Data is stored at cell centers Information is needed at cell faces. APPROACH i+1/2,j,k L R i-1, j, k i, j, k i+1, j, k

  17. Let us approximate qi+1/2 in the uniform transformed plane with three points: i+1/2 i+1 i-1 i APPROACH

  18. Using classical Taylor series method, we can obtain three expansion equations of qi+1, qi, and qi-1 about i+1/2, for example: With these three equations, we can determine coefficients ai+1, ai, and ai-1 (Traditional Method). APPROACH

  19. In our approach, we impose a further restriction to match the Fourier transformation (in space) of approximation for qi+1/2with its exact transformation. The Fourier transformation of approximate expression for qi+1/2 is: APPROACH F.T.

  20. The following error expression should be minimized: with respect to coefficients . This leads to an over-determined system. Solved by Least Square method. APPROACH

  21. Standard 3rd Order Monotone Upstream-centered Scheme for the Conservative Law (MUSCL Scheme): Present Scheme: APPROACH

  22. High-Speed Impulsive noise modeling Preliminary studies of Blade-Vortex Interaction noise Tip vortex system prediction RESULTS

  23. 1/7 scale model of untwisted rectangular UH-1H blades in hover condition NACA0012 airfoil Non-lifting case Shock Noise Test Parameters

  24. Microphone Shock wave R Blade r/R=1.78 r/R=1.111 Shock Noise Measurement Locations and Method

  25. Variation of Acoustic Pressure p´ with time for a Non-lifting Rotor, MTip= 0.88, r/R=1.136, Grid size 1335535

  26. Variation of Acoustic Pressure p´ with time for a Non-lifting Rotor, MTip= 0.88, r/R=3.09, Grid size 1335535

  27. Variation of Acoustic Pressure p´ with time for a Non-lifting Rotor, MTip = 0.9, r/R=1.111, Grid size 1335535

  28. Variation of Acoustic Pressure p´ with time for a Non-lifting Rotor, MTip = 0.9,r/R=3.09, Grid size 1335535

  29. Variation of Acoustic Pressure p´ with time for a Non-lifting Rotor, MTip= 0.95, r/R=1.053, Grid size 1335535

  30. Variation of Acoustic Pressure p´ with time for a Non-lifting Rotor, MTip= 0.95,r/R=3.09, Grid size 1335535

  31. BLADE-VORTEX PROXIMITY Zv Y X VORTEX GENERATOR NEAR FIELD MICROPHONES Z +CCW VORTEX ROTATION V +v Parallel BVI Study Schematic of experimental set-up in wind tunnel test section

  32. Untwisted, rectangular blade NACA 0012 airfoil Mtip=0.71, Advance ratio=0.2 Vortex 0.25 chord below blade Non-lifting case Parallel BVI Test Parameters

  33. Parallel BVI Study(169  45  57) Near-field acoustic pressure for microphone 7

  34. 1/7 scale model of Operational Load Survey (OLS) blades Rectangular blades with 8.2 of twist from root to tip Mtip=0.664, Advance ratio=0.164 Grid size 110  45  40 AH-1 Forward Flight Test Parameters

  35. AH-1 Forward Flight Self-induced wake Descending direction Interaction of tip vortices with rotor disk in descending flight

  36. =90 Advancing Side     Inlet Flow =0 =180  Tip Vortex Retreating Side  AH-1 Forward Flight Schematic of flow field

  37. AH-1 Forward Flight Blade Surface Pressure Coefficient Distribution, r/R=0.955, =0

  38. AH-1 Forward Flight Blade Surface Pressure Coefficient Distribution, r/R=0.955, =90

  39. AH-1 Forward Flight Blade Surface Pressure Coefficient Distribution, r/R=0.955, =180

  40. How well does the Low Dispersion Scheme model tip vortices? Schematic of hover rotor wake structure

  41. Untwisted rectangular NACA0012 blades Hovering condition MTip=0.44 Collective Pitch c=8 Caradonna & Tung Rotor Test Parameters

  42. TURNS-MUSCL TURNS-LDFV Vortex I Vortex I Vortex II Vorticity Magnitude Contour Caradonna & Tung RotorMTip=0.44

  43. Caradonna & Tung RotorMTip=0.44, r/R=0.80, Grid size 79  45  31 Blade Surface Pressure Distribution

  44. A Low-Dispersion Finite Volume scheme has been developed and implemented into TURNS, a finite volume CFD code. Encouraging agreement between the predicted results and experiment data has been obtained for shock noise on coarse grid. CONCLUSIONS

  45. Basic characteristics of BVI noise are predicted with satisfactory accuracy. TURNS-LDFV can capture main features of the tip vortex system with good resolution on coarse grids. CONCLUSIONS

  46. Determine the minimum number of grid points needed to predict shock noise. Identify the contributions of different noise sources. PROPOSED WORK

  47. Repeat forward flight BVI calculation on fine grid; Incorporate trim effects. Further investigation of BVI noise investigated in Higher harmonic control Aeroacoustic Rotor Test (HART) program. PROPOSED WORK

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