Gary G. Podboy

Steady and Unsteady Flow Field Measurements Within a NASA 22-Inch Fan Model

Results are presented of an experiment conducted to investigate possible sources of fan noise in the flow developed by a 22-in. (55.9 cm) diameter turbofan model. Flow diagnostic data were acquired to identify possible sources of both tone and broadband noise. Laser Doppler velocimetry was used to characterize the tip flows that develop within the rotor blade passages, the wake flow downstream of the rotor, and the shock waves that develop on the blades when operated at transonic relative tip speeds. Single-point hot-wire measurements were made in the rotor wake to determine the frequency content and the length scales of the flow unsteadiness. The results document the changes in the rotor wake flow with both rotor speed and axial distance downstream of the rotor. The data also show the tip flow development within the blade passage, its migration downstream, and (at high rotor speeds) its merging with the blade wake of the following blade. Data also depict the variation of the tip flow with tip clearance. LDV data obtained within the blade passages at high rotor speeds illustrate the passage-to-passage variation of the mean shock position. Spectra computed from the single-point hot-wire measurements illustrate how the energy in the flow oscillations is split between periodic and random components, and how this split varies with both radial and axial position in the rotor wake.

Improved Phased Array Imaging of a Model Jet

Abstract An advanced phased array system, OptiNav Array 48, and a new deconvolution algorithm, TIDY, have been used to make octave band images of supersonic and subsonic jet noise produced by the NASA Glenn Small Hot Jet Rig (SHJAR). The results are much more detailed than previous jet noise images. Shock cell structures and the production of screech in an underexpanded supersonic jet are observed directly. Some trends are similar to observations using spherical and elliptic mirrors that partially informed the two-source model of jet noise, but the radial distribution of high frequency noise near the nozzle appears to differ from expectations of this model. The beamforming approach has been validated by agreement between the integrated image results and the conventional microphone data. Introduction Jet noise remains interesting after many years of study, despite a complete lack of hardware in the noise-producing region. The simplest case, a single-flow round jet with no forward flight, can produce a range of noise source phenomena: turbulent shear flow, mixing noise from large scale coherent structures, broadband shock associated noise, and screech tones (Ref. 1). These have been studied by measuring directivity in the far field, the acoustic pressure distribution in the near field, microphone two-point correlation, and spherical and elliptic mirror imaging (Ref. 2). Shock cells occur in the supersonic case when the jet is not ideally expanded. Flow disturbances convecting through the shock cells give rise to broadband shock-associated noise and screech tones (Refs. 3 and 4). Shock-associated noise can be identified in far-field spectra as an excess above the mixing noise that follows a well-defined correlation (Ref. 5). The shock cell structure can be imaged with Schlieren photography and its fluctuations have been studied with optical deflectrometry (Ref. 6) and numerical methods (Refs. 7 and 8). Several studies of jet noise using phased arrays and related systems of microphone arrays have been made (Refs. 9 to 18). These measurements have frequently been limited to subsonic jets, and have had several drawbacks that prevent the structure of the jet noise source from being fully revealed. One-dimensional arrays (Refs. 9, 14, and 18) do not readily show the radial structure of the source. Cage arrays (Refs. 10, 12, and 15) are designed for three-dimensional imaging, but they have several drawbacks. They subtend an angle that is too large to illustrate the source for different directivity angles. Their location in the near field can give confusing results. Finally, the large separation of the microphones produces severe problems with turbulent decorrelation. Phased array studies of jet noise have been criticized on the basis that the monopole source model used in beamforming is too simplistic. This issue has been addressed by Michel (Ref. 19), who showed that jet source convection and correlation length effects can be reduced to simple sources with directivity patterns. The use of a phased array is therefore justified, provided that it subtends a reasonably small range of solid angle. A similar argument was made to justify mirror studies of jets in the 1970s (Ref. 20). A straightforward but conservative specification would be to require that the array is small enough that its Rayleigh resolution spot size is larger than the presumed correlation length of the turbulence. The Rayleigh resolution of imaging system of this type (essentially a microscope) is

Jet-Surface Interaction Test: Phased Array Noise Source Localization Results

An experiment was conducted to investigate the effect that a planar surface located near a jet flow has on the noise radiated to the far-field. Two different configurations were tested: 1) a shielding configuration in which the surface was located between the jet and the far-field microphones, and 2) a reflecting configuration in which the surface was mounted on the opposite side of the jet, and thus the jet noise was free to reflect off the surface toward the microphones. Both conventional far-field microphone and phased array noise source localization measurements were obtained. This paper discusses phased array results, while a companion paper discusses far-field results. The phased array data show that the axial distribution of noise sources in a jet can vary greatly depending on the jet operating condition and suggests that it would first be necessary to know or be able to predict this distribution in order to be able to predict the amount of noise reduction to expect from a given shielding configuration. The data obtained on both subsonic and supersonic jets show that the noise sources associated with a given frequency of noise tend to move downstream, and therefore, would become more difficult to shield, as jet Mach number increases. The noise source localization data obtained on cold, shock-containing jets suggests that the constructive interference of sound waves that produces noise at a given frequency within a broadband shock noise hump comes primarily from a small number of shocks, rather than from all the shocks at the same time. The reflecting configuration data illustrates that the law of reflection must be satisfied in order for jet noise to reflect off of a surface to an observer, and depending on the relative locations of the jet, the surface, and the observer, only some of the jet noise sources may satisfy this requirement.

Limitations of Phased Array Beamforming in Open Rotor Noise Source Imaging

Phased array beamforming results of the F31/A31 historical baseline counter-rotating open rotor blade set were investigated for measurement data taken on the General Electric Counter-Rotating Open Rotor Propulsion Rig in the 9’ x 15’ Low Speed Wind Tunnel of NASA Glenn Research Center as well as data produced using the LINPROP open rotor tone noise code. The planar microphone array was positioned broadside and parallel to the axis of the open rotor, roughly 2.3 rotor diameters away. The results provide insight as to why the apparent noise sources of the blade passing frequency tones and interaction tones appear at their nominal Mach radii instead of at the actual noise sources, even if those locations are not on the blades. Contour maps corresponding to the sound fields produced by the radiating sound waves, taken from the simulations, are used to illustrate how the interaction patterns of circumferential spinning modes of rotating coherent noise sources interact with the phased array, often giving misleading results, as the apparent sources do not always show where the actual noise sources are located. This suggests that a more sophisticated source model would be required to accurately locate the sources of each tone. The results of this study also have implications with regard to the shielding of open rotor sources by airframe empennages.

Effect of Trailing Edge Flow Injection on Fan Noise and Aerodynamic Performance

An experimental investigation using trailing edge blowing for reducing fan rotor/guide vane wake interaction noise was completed in the NASA Glenn 9- by 15-foot Low Speed Wind Tunnel. Data were acquired to measure noise, aerodynamic performance, and flow features for a 22” tip diameter fan representative of modern turbofan technology. The fan was designed to use trailing edge blowing to reduce the fan blade wake momentum deficit. The test objective was to quantify noise reductions, measure impacts on fan aerodynamic performance, and document the flow field using hot-film anemometry. Measurements concentrated on approach, cutback, and takeoff rotational speeds as those are the primary conditions of acoustic interest. Data are presented for a 2% (relative to overall fan flow) trailing edge injection rate and show a 2 dB reduction in Overall Sound Power Level (OAPWL) at all fan test speeds. The reduction in broadband noise is nearly constant and is approximately 1.5 dB up to 20 kHz at all fan speeds. Measurements of tone noise show significant variation, as evidenced by reductions of up to 6 dB in the 2 BPF tone at 6700 rpmc and increases of nearly 2 dB for the 4 BPF tone at approach speed. Aerodynamic performance measurements show the fan with 2% injection has an overall efficiency that is comparable to the baseline fan and operates, as intended, with nearly the same pressure ratio and mass flow parameters. Hot-film measurements obtained at the approach operating condition indicate that mean blade wake filling in the tip region was not as significant as expected. This suggests that additional acoustic benefits could be realized if the trailing edge blowing could be modified to provide better filling of the wake momentum deficit. Nevertheless, the hot-film measurements indicate that the trailing edge blowing provided significant reductions in blade wake turbulence. Overall, these results indicate that further work may be required to fully understand the proper implementation of injecting flow at/near the trailing edge as a wake filling strategy. However, data do support the notion that noise reductions can be realized not only for tones but perhaps more importantly, also for broadband. Furthermore, the technique can be implemented without adversely effecting overall fan aerodynamic performance.

Noise Benefits of Rotor Trailing Edge Blowing for a Model Turbofan

*† ‡ An advanced model turbofan was tested in the NASA Glenn 9- by 15-Foot Low Speed Wind Tunnel (9x15 LSWT) to explore far field acoustic effects associated with rotor Trailing-Edge-Blowing (TEB) for a modern, 1.294 stage pressure ratio turbofan model. The TEB rotor (Fan9) was designed to be aerodynamically similar to the previously tested Fan1, and used the same stator and nacelle hardware. Fan9 was designed with trailing edge blowing slots using an external air supply directed through the rotor hub. The TEB flow was heated to approximate the average fan exit temperature at each fan test speed. Rotor root blockage inserts were used to block TEB to all but the outer 40 and 20% span in addition to full-span blowing. A configuration with full-span TEB on alternate rotor blades was also tested. Far field acoustic data were taken at takeoff/approach conditions at 0.10 tunnel Mach. Far-field acoustic results showed that full-span blowing near 2.0% of the total flow could reduce the overall sound power level by about 2 dB. This noise reduction was observed in both the rotor-stator interaction tones and for the spectral broadband noise levels. Blowing only the outer span region was not very effective for lowering noise, and actually increased the far field noise level in some instances. Full-span blowing of alternate blades at 1.0% of the overall flow rate (equivalent to full-span blowing of all blades at 2.0% flow) showed a more modest noise decrease relative to full-span blowing of all blades. Detailed hot film measurements of the TEB rotor wake at 2.0% flow showed that TEB was not every effective for filling in the wake defect at approach fan speed toward the tip region, but did result in overfilling the wake toward the hub. Downstream turbulence measurements supported this finding, ∗ and support the observed reduction in spectral broadband noise. I. Introduction n advanced model turbofan was tested in the NASA Glenn 9x15 LSWT to explore far field acoustic effects associated with rotor TEB. Previous research has shown that the rotor viscous wake may be significantly reduced by injecting “make up” air at or near the trailing edge. Reduction of the rotor wake results in less wake interaction with the downstream stator and therefore less rotor-stator interaction noise. This noise reduction is particularly beneficial for the interaction tone levels at multiples of Blade Passing Frequency (nBPF), but may also be seen in the broadband noise levels. This paper presents far field acoustic results for a modern high bypass ratio, low tip speed model turbofan with rotor trailing edge blowing. These model tests were conducted in an anechoic low-speed wind tunnel at the NASA Glenn Research Center in 2005. The concept of reducing blade wake through trailing edge blowing has been investigated in the literature. Reference 1 shows water tunnel results for a blade with trailing edge jets. This reference showed that the blade wake defect could be significantly reduced by proper injection of makeup fluid at the blade trailing edge, and suggests that this technique might have benefits for turbofan engines. Brookfield and Waitz 2 tested a modern 22-in. diameter model turbofan with rotor TEB in a transient blow-down facility. Their results showed that the first three BPF harmonics amplitudes could be reduced by as much as 85% (measured in-duct 1.5 chords downstream of the rotor) using a blowing mass flow less than 2% of the stage throughflow. Although encouraging, these results were compromised by facility limitations, including non-anechoic accommodations for far field noise measurements.

Measurements of two-point velocity correlations in a round jet with application to jet noise

This work documents two-point space-time correlation measurements made in a cold, M=OS round jet using a pair of two-component hotwire probes, From the simultaneously measured velocities space-time correlations are computed for axial and radial displacements for several radial locations each at three axial stations (4, 8, and 20 diameters). This rich type of turbulence data is required in many jet noise prediction theories, such as the Mani-Gliebe-Balsa (MGB) theory. In the past, these terms have been estimated from an isotropic, homogeneous turbulence model with a Gaussian form for space and time dependence. This turbulence model resulted in a particular partitioning of turbulent kinetic energy among the many quadrupole components. The work presented here indicates that such models are in error and proposes an axisymmetric model with lower order exponent. This model results in a different ratio of source directivities, which will impact the predicted jet noise source directivity. Using a lower-order exponent in the time-dependent part of the space-time correlation also shows better agreement with data and will change the predicted jet noise spectral characteristics.

Phased Array Noise Source Localization Measurements of an F404 Nozzle Plume at Both Full and Model Scale

A 48-microphone planar phased array system was used to acquire jet noise source localization data on both a full-scale F404-GE-F400 engine and on a 1/4th scale model of a F400 series nozzle. The full-scale engine test data show the location of the dominant noise sources in the jet plume as a function of frequency for the engine in both baseline (no chevron) and chevron configurations. Data are presented for the engine operating both with and without afterburners. Based on lessons learned during this test, a set of recommendations are provided regarding how the phased array measurement system could be modified in order to obtain more useful acoustic source localization data on high-performance military engines in the future. The data obtained on the 1/4th scale F400 series nozzle provide useful insights regarding the full-scale engine jet noise source mechanisms, and document some of the differences associated with testing at model-scale versus full-scale.

Acoustic Performance of Novel Fan Noise Reduction Technologies for a High Bypass Model Turbofan at Simulated Flights Conditions

Two novel fan noise reduction technologies, over the rotor acoustic treatment and soft stator vane technologies, were tested in an ultra-high bypass ratio turbofan model in the NASA Glenn Research Center’s 9x15 Low Speed Wind Tunnel. The performance of these technologies was compared to that of the baseline fan configuration, which did not have these technologies. Sideline acoustic data and hot film flow data were acquired and are used to determine the effectiveness of the various treatments. The material used for the over the rotor treatment was foam metal and two different types were used. The soft stator vanes had several internal cavities tuned to target certain frequencies. In order to accommodate the cavities it was necessary to use a cut-on stator to demonstrate the soft vane concept.

Aeroacoustic Analysis of Fan Noise Reduction With Increased Bypass Nozzle Area

An advanced model turbofan was tested in the NASA Glenn 9-by 15-Foot Low Speed Wind Tunnel (9x15 LSWT) to explore far field acoustic effects of increased bypass nozzle area. This fan stage test was part of the NASA Glenn Fan Broadband Source Diagnostic Test, second entry (SDT2) which acquired aeroacoustic results over a range of test conditions. The baseline nozzle was sized to produce maximum stage performance at cruise condition. However, the wind tunnel testing is conducted near sea level condition. Therefore, in order to simulate and obtain performance at other operating conditions, two additional nozzles were designed and tested one with +5 percent increase in weight flow (+5.4 percent increase in nozzle area compared with the baseline nozzle), sized to simulate the performance at the stage design point (takeoff) condition, and the other with a +7.5 percent increase in weight flow (+10.9 percent increase in nozzle area) sized for maximum weight flow with a fixed nozzle at sea level condition. Measured acoustic benefits with increased nozzle area were very encouraging, showing overall sound power level (OAPWL) reductions of 2 or more dB while the stage thrust actually increased by 2 to 3 percent except for the most open nozzle at takeoff rotor speed where stage performance decreased. Effective perceived noise levels for a 1500 ft engine flyover and 3.35 scale factor showed a similar noise reduction of 2 or more EPNdB. Noise reductions, principally in the level of broadband noise, were observed everywhere in the far field. Laser Doppler Velocimetry measurements taken downstream of the rotor showed that the total turbulent velocity decreased with increasing nozzle flow, which may explain the reduced rotor broadband noise levels.