Evaluating the Acoustic Effect of Over-the-Rotor Foam-Metal Liner Installed on a Low Speed Fan Using Virtual Rotating Microphone Imaging

An in-duct beamforming technique for imaging rotating broadband fan sources has been used to evaluate the acoustic characteristics of a Foam-Metal Liner installed over-the-rotor of a low-speed fan. The NASA Glenn Research Center s Advanced Noise Control Fan was used as a test bed. A duct wall-mounted phased array consisting of several rings of microphones was employed. The data are mathematically resampled in the fan rotating reference frame and subsequently used in a conventional beamforming technique. The steering vectors for the beamforming technique are derived from annular duct modes, so that effects of reflections from the duct walls are reduced

Aeroacoustics of Propulsion Airframe Integration: Overview of NASA’s Research,” Noise Con Paper 2003105, presented at the 2003 Noise Con Conference

The integration of propulsion and airframe is fundamental to the design of an aircraft system. Many considerations influence the integration, such as structural, aerodynamic, and maintenance factors. In regard to the acoustics of an aircraft, the integration can have significant effects on the net radiated noise. Whether an engine is mounted above a wing or below can have a significant effect on noise that reaches communities below because of shielding or reflection of engine noise. This is an obvious example of the acoustic effects of propulsion airframe installation. Another example could be the effect of the pylon on the development of the exhaust plume and on the resulting jet noise. In addition, for effective system noise reduction the impact that installation has on noise reduction devices developed on isolated components must be understood. In the future, a focus on the aerodynamic and acoustic interaction effects of installation, propulsion airframe aeroacoustics, will become more important as noise reduction targets become more difficult to achieve. In addition to continued fundamental component reduction efforts, a system level approach that includes propulsion airframe aeroacoustics will be required in order to achieve the 20 dB of perceived noise reduction envisioned by the long-range NASA goals. This emphasis on the aeroacoustics of propulsion airframe integration is a new part of NASA’s noise research. The following paper will review current efforts and highlight technical challenges and approaches. 1

CONTROL OF INFLOW DISTORTION IN A SCARF INLET

The scarf inlet has the potential to reduce aircraft inlet noise radiation to the ground by reflecting it into the space above the engine. Without forward motion of the engine, the non-symmetry of the inlet causes inflow distortion which generates noise that is greater than the noise reduction of the scarf. However, acoustic evaluations of aircraft engines are often done on static test stands. A method to reduce inflow distortion by boundary layer suction is proposed and evaluated using a model of a high bypass ratio engine located in an anechoic chamber. The design goal of the flow control system is to make the inflow to the inlet circumferentially uniform and to eliminate reversed flow. This minimizes the inflow distortion and allows for acoustic evaluation of the scarf inlet on a static test stand. The inlet boundary layer suction effectiveness is evaluated both by aerodynamic and by acoustic measurements. Although the design goal is not met, the control system is found to have a beneficial effect on the engine operation, reducing blade stall and speed variation. This is quantified by two acoustic benefits, reduction both of the variability of tone noise and of the low frequency wideband noise due to the inflow distortion. It is felt that a compromise in the manufacture of the control hardware contributes to the inability of the control system to perform as expected from the analysis. The control system with sufficient authority is felt to have the potentia

Abstract

The NASA Aeroacoustic Prediction System (NAPS) is used to establish a link between model-scale and full-scale rotor predictions and is partially validated against measured wind tunnel and flight aeroacoustic data. The prediction approach of NAPS couples a comprehensive rotorcraft analysis with acoustic source noise and propagation codes. The comprehensive analysis selected for this study is CAMRAD-II, which provides the performance/trim/wake solution for a given rotor or flight condition. The post-trim capabilities of CAMRAD-II are used to compute high-resolution sectional airloads for the acoustic tone noise analysis, WOPMOD. The tone noise is propagated to observers on the ground with the propagation code, RNM (Rotor Noise Model). Aeroacoustic predictions are made with NAPS for an isolated rotor and compared to results of the second Harmonic Aeroacoustic Rotor Test (HART-II) program, which tested a 40 % dynamically and Mach-scaled BO-105 main rotor at the DNW. The NAPS is validated with comparisons for three rotor conditions: a baseline condition and two Higher Harmonic Control (HHC) conditions. To establish a link between model and full-scale rotor predictions, a full-scale BO-105 main rotor input deck for NAPS is created from the 40 % scale rotor input deck. The full-scale isolated rotor predictions are then compared to the model predictions. The comparisons include aerodynamic loading, acoustic levels, and acoustic pressure time histories for each of the three conditions. With this link established, full-scale predictions are made for a range of descent flight conditions and compared with measured trends from the recent Rotorcraft Operational Noise Abatement Procedures (RONAP) flight test conducted by DLR and ONERA. Additionally, the effectiveness of two HHC conditions from the HART-II program is demonstrated for the full-scale rotor in flight. BPF Notation Blade passage frequenc

Abstract Analysis of Measured and Predicted Acoustics

Flight acoustic and vehicle state data from an XV-15 acoustic flight test are examined. Flight predictions using TRAC are performed for a level flight (repeated) and four descent conditions (including a BVI). The assumptions and procedures used for TRAC flight predictions as well as the variability in flight measurements, which are used for input and comparison to predictions, are investigated in detail. Differences were found in the measured vehicle airspeed, altitude, glideslope, and vehicle orientation (yaw, pitch and roll angle) between each of the repeat runs. These differences violate some of the prediction assumptions and significantly impacted the resulting acoustic predictions. Multiple acoustic pulses, with a variable time between the pulses, were found in the measured acoustic time histories for the repeat runs. These differences could be attributed in part to the variability in vehicle orientation. Acoustic predictions that used the measured vehicle orientation for the repeat runs captured this multiple pulse variability. Thickness noise was found to be dominant on approach for all the cases, except the BVI condition. After the aircraft passed overhead, broadband noise and low frequency loading noise were dominant. The predicted LowSPL time histories compared well with measurement on approach to the array for the non-BVI conditions and poorly for the BVI condition

The Devil is in the Details

The accurate prediction of the aeroacoustic eld generated by aerospace vehicles or nonaerospace machinery is necessary for designers to control and reduce source noise. Powerful computational aeroacoustic methods, based on various acoustic analogies (primarily the Lighthill acoustic analogy) and Kirchho methods, have been developed for prediction of noise from complicated sources, such as rotating blades. Both methods ultimately predict the noise through a numerical evaluation of an integral formulation. In this paper, we consider three generic acoustic formulations and several numerical algorithms that have been used to compute the solutions to these formulations. Algorithms for retarded-time formulations are the most e cient and robust, but they are di cult to implement for supersonic-source motion. Collapsing-sphere and emission-surface formulations are good alternatives when supersonicsource motion is present, but the numerical implementations of these formulations are more computationally demanding. New algorithms|which utilize solution adaptation to provide a speci ed error level|are needed. Notation 2 wave operator, D'Alembertian c sound speed in undisturbed medium d6 element of emission (in uence) surface d element of collapsing-sphere surface dS element of source surface f = 0 function that describes the source surface g = 0 surface that describes the collapsing sphere, g = 0 t + r=

Sensitivity analysis and uncertainty quantification for the ffowcs williams-hawkings equation

The acoustic propagation stage of a computational aeroacoustic analysis has been in- vestigated for possible sources of uncertainty and sensitivity. The acoustic propagation is realized through an acoustic prediction module that is fundamentally based on the Ffowcs Williams-Hawkings equation. Non-intrusive polynomial chaos expansion methods are used, along with a direct sensitivity analysis based on Sobol indices. Three analytical test cases are chosen in order to isolate the acoustic propagation stage from the noise source identification stage. In an analysis of a theoretical helicopter blade, it is identified that the mean flow Mach number and blade tip Mach number are significant contributors to noise uncertainty. As the advancing blade tip Mach number approaches the transonic and supersonic flow regimes, these uncertainties are amplified. The source of the uncertainty is mainly attributed to the blade tip Mach number at low mean flow Mach numbers, however as the mean flow Mach number increases, the contribution of the mean flow Mach number to the uncertainty significantly increases. © 2017, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved

Acoustic Characterization of a Helmholtz Resonator Under Grazing Flow Conditions Using a Hybrid Methodology

A hybrid methodology is applied to characterize a Helmholtz resonator under a grazing flow. In a first step, the mean flow profile is obtained from a RANS simulation. The acoustic field and its interaction with the hydrodynamic field are calculated in a second step, by solving the linearized Navier-Stokes equations. The impedance of the Helmholtz resonator is obtained using the in-situ measurement technique. Because the method relies on the use of linear governing equations, the validity of the results is limited to the linear regime. Within this region, the computed reactance shows good agreement with an analytical model and experimental observations. Preliminary computations of the resistance show the potential of the method, but the optimal position of the facing sheet measurement point remains an open question