jet aircraft propulsion noise reduction

American Institute of Aeronautics and Astronautics

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Jet Aircraft Propulsion Noise Reduction Research at University of Cincinnati

Olaf Rask*, Seth Harrison†, David Munday‡, Chris Harris§, Dr. Mihai Mihaescu** and Dr. Ephraim Gutmark†† University of Cincinnati, Cincinnati, Ohio, 45221

The University of Cincinnati has an active Aeroacoustics program studying the application of several flow control technologies to typical separate-flow exhaust systems like those found on modern jet engines. The University of Cincinnati Aeroacoustic Test Facility (UC-ATF) includes a high fidelity model of such an exhaust system with interchangeable hardware allowing the simulation of medium- and high-bypass geometries and the application of various nozzles with different technologies to reduce the generated noise. Various chevron geometries, steady and pulsed blowing, and other approaches are tested and evaluated. Those technologies which show merit are then selected for further test in collaboration with our industrial partners at GE Aviation and GE Global Research. Testing at GE allows larger scale and improved fidelity. This testing pipeline allows further down-selection path to full-scale engine tests for the best technologies. In this way UC contributes to the missions of our industrial partners while advancing the state of knowledge regarding jet noise reduction technologies.

Nomenclature Deq = Equivalent diameter M = Mach number NPR = Nozzle pressure ratio OASPL = Overall sound pressure level SPL = Sound pressure level St = Strouhal number

I. Introduction IRCRAFT noise is a significant concern to those who live near any large airport. As the volume of air traffic increases, so does the impact on those increasing numbers who live near our busy airports. The regulatory

response to limiting aircraft noise is embodied in Federal Aviation Regulation (FAR) Chapter 36 in the Unites States and in International Civil Aeronautics Organization (ICAO) Annex 16 elsewhere which impose limits which become increasingly stringent with time. Manufacturers of aircraft and aircraft engines face the technical challenge of making aircraft simultaneously quieter, more powerful, and more efficient. The conflicting requirements of these goals motivate the present research, which seeks to apply flow control to one source of aircraft noise in a manner that will yield acoustic benefit while minimizing the impact on performance.1

One of the more significant sources of aircraft noise in modern jet aircraft is the turbulence generated in the shear layers around the engine’s exhaust. For the separate flow exhaust designs common on large commercial aircraft there are two such shear layers generating noise, the inner and outer shear layers. The inner shear layer is the layer between the primary, or core flow, and the secondary, or fan flow. The outer shear layer lies between the secondary flow and the freestream. At any useful operating condition we will have a significant shear velocity across one or both of these shear layers. These shear layers are unstable and the instabilities lead to vortex rollup

* Graduate Student, Aerospace Engineering, [email protected], AIAA Student Member. † Graduate Student, Aerospace Engineering, [email protected], AIAA Student Member. ‡ Graduate Student, Aerospace Engineering, [email protected], AIAA Student Member. § Graduate Student, Aerospace Engineering, [email protected], AIAA Student Member. ** Research Assistant Professor, Aerospace Engineering, [email protected], AIAA Member. †† Professor, Ohio Eminent Scholar, Aerospace Engineering, [email protected], AIAA Associate Fellow.

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and transition to turbulence. The coherent structures and turbulent eddies generate non-equilibrium pressure fluctuations which are radiated as sound.2-6

A number of flow control approaches have been applied to modify the flow structures in the shear layer and the radiated sound. One of the simplest and widely accepted approaches which has begun entering service on the commercial fleet is the application of chevrons to the trailing edge of the nozzles. Unlike conventional, round nozzles, chevron nozzles employ serrations on the trailing edge. These chevrons typically impinge slightly into the higher-speed flow and cause the flow passing over them to turn and curl around their trailing edges, thus introducing a rotating component to the velocity field. The net effect of each chevron is to shed a pair of streamwise vortices. These vortices speed the mixing between the flows and bring them more quickly to a low-shear condition. This has been shown to reduce the overall level of noise produced. It typically reduces the sound pressure level (SPL) at low frequencies at the expense of increasing the levels at higher frequencies. A disadvantage of chevrons is that they impinge into the flow and produce a reduction in thrust. This thrust loss is an acceptable trade at take-off, but at cruise, where the need for noise reduction is less, the cost is less justified.7-9

An alterative which has shown promise is to introduce similar vortical motion into the shear layers by directly blowing air into the shear layers at an angle to the main flow. Pairs of steady blowing jets can create counter-rotating vortex pairs just as chevrons do. A significant advantage of such blowing is that it can be turned off when not needed. The bleed air required to drive the small jets introduces an undoubted performance penalty, but once the aircraft has left the noise-sensitive airport environment, the bleed can be turned off and the penalty is not incurred at cruise.10 This paper serves as an overview of the experimental and computational work by the UC Aeroacoustics team.

II. Experimental Facility and Procedure The model under test is a subscale model of a typical commercial jet engine with separate flow exhaust of

medium bypass. (Figure 1a) The model is mounted in the University of Cincinnati Acoustic Test Facility (UC-ATF) which is a 24’ x 25’ x 11’ test chamber which has been acoustically treated to be anechoic down to 500 Hz. Figure 1b shows the layout of the facility. For far-field measurements eight quarter-inch microphones are arrayed along the depicted arc at angles from 70° to 150° measured from the upstream direction. The microphones were placed at 35 equivalent diameters from the center of the fan exit. The model and facility are more fully described in by Callender in Reference 11. The microphone signals are filtered above 100 kHz and recorded at 200 kHz for 10 seconds. The flow variables are monitored at 4 Hz in order to establish and hold the desired operating condition of the model. Variables recorded include the ambient conditions in the chamber as well as the fan and core flow stagnation pressures and core stagnation temperature.

The model under test allows quick and easy installation of interchangeable hardware such as the chevron nozzles in Figure 2, or blowing hardware as in Figure 3a.

III. Results and Discussion

A. Subsonic Chevrons As a noise reduction technique, chevron nozzles are both practical and mature enough that they are now in

revenue generating service. These nozzles differ from traditional, conical nozzles in that their trailing edges are not smooth, but have triangular cutouts. With a significant velocity shear, the difference in dynamic pressure forces the high speed flow to wrap around each edge of individual chevrons. This wrapping introduces rotation, forming pairs of counter-rotating vortices behind each chevron. These vortices enhance mixing between layers and reduce velocity shear more quickly. This has been shown to reduce the noise produced by mixing. However, as the vortices move downstream, they become unstable and break down. This vortex breakdown is an additional source of fine scale turbulence and high frequency noise.

The two principal parameters that effect noise reduction are shear velocity and penetration. When either of these are increased, their effects also increase. Increasing shear velocity or penetration will always reduce low frequency noise, but may cause an increase in high-frequency noise. These effects can be balanced in order to achieve a net reduction in the integrated overall sound pressure level (OASPL). Reductions in OASPL of as much as 2-5 dB have been achieved using chevrons on both the core and fan nozzles. 12

In exploring the interaction between core and fan chevrons, tests were run at a number of cycle conditions chosen to emphasize different aspects of the flow. Baseline conical nozzles were compared to chevron nozzles on both streams. On the core stream an eight-lobed low penetration chevron nozzle was used. On the fan stream a


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sixteen-lobed high-penetration chevron nozzle was used. We chose to test these nozzles together because they had both shown significant acoustic benefit separately.

Cycle conditions were chosen to isolate the effects of each chevron nozzle and to examine the interaction between them. One condition was chosen in which the core/fan shear velocity was negligible in order to isolate the fan/ambient shear layer effects. A second condition was chosen with negligible fan/ambient shear velocity to isolate core/fan shear layer effects. Finally a condition was chosen in which both shear velocities were significant. For this condition the velocities were normalized by the mass-averaged velocity. The shear velocity between the core and fan stream was 0.84 the shear between the fan stream and quiescent ambient air was 0.55. In all cases the core NPR was 1.85 and the core temperature was 220F. The fan temperature was at 80F.

Far-field results for the third condition are shown in Figure 4a. Two directivity angles are shown, 90° and 150°. At this cycle point, there is an appreciable shear across both nozzles, so it is expected to see acoustic effects from both fan and chevron nozzles.

We first look at the side-line data. Examination of the spectra at 90º directivity angle reveals a 1dB low-Strouhal-number reduction when using only fan chevrons. Using only core chevrons we see a slightly larger reduction of 2dB at low Strouhal number and a slight increase above a Strouhal number of 4. Fan and core chevrons together result in a maximum 3dB reduction at low Strouhal number, with a 1 dB penalty at high Strouhal number.

Now looking at the aft-angle data, the 150º directivity, shows that the use of fan chevrons alone results in a uniform 1 dB reduction at all Strouhal numbers. Use of core chevrons alone results in a maximum 4 dB reduction at low Strouhal number and no increase at any Strouhal number. Use of fan and core chevrons together results in a maximum 6dB low-Strouhal-number reduction and a 2 dB high-Strouhal-number reduction. At all Strouhal numbers for this angle, fan and core chevrons reduce noise more than any other nozzle combination.

We now look at two single Strouhal numbers representative of the high-frequency and low-frequency regions. In Figure 4b we examine the low-Strouhal-number reduction at the aft angle of 150°, and the high-Strouhal-number increase at the side angle of 90°. Modest low-Strouhal-number reductions of 1dB are demonstrated when using only fan chevrons. Significant reductions of 4.3 dB can be seen when using only core chevrons. However, the best reductions of 5.5 dB are realized when using fan and core chevrons at the same time. The sum of the reductions due to fan and core chevrons alone is roughly equal to the reduction when using both chevron nozzles together.

The noise increase at high Strouhal number follows a similar trend. The core chevron nozzle alone produced an increase of 0.1dB. This is within the measurement uncertainty and is thus negligible. Using fan chevrons alone produced a somewhat larger increase of 0.4dB. High-Strouhal-number noise increase was greatest when using fan and core chevrons together. This configuration produced an increase of 1dB.

The OASPL data is shown in Figure 5. There are modest reductions to jet noise when using just fan chevrons. There are significantly better reductions when using only core chevrons. However, it is apparent that reductions realized when using fan and core chevrons are far superior to reductions seen when they are used independently. The reductions due to fan and core chevrons together appear to be roughly equal to the sum of reductions due to only fan chevrons plus the reductions due to only core chevrons. These two values are within 5% of each other at every angle. This suggests that the effects of the two chevrons may be independent of one another.

Near-field results at select Strouhal numbers are shown in Figure 6.13 The chevron configuration is shown by the hardware cartoon in each subfigure. Column (a) shows a Strouhal number of 0.45. This is the frequency at which the greatest low-frequency reductions were evident in the far-field spectra. The baseline case shows two regions of noise generation. The first, between one and five equivalent diameters is the region of sound generation from the fan/ambient shear layer. The second region is between six and ten equivalent diameters. This region is associated with noise production by the core/fan shear layer. Application of core chevrons reduces this noise, supporting the thesis. Finally application of both fan and core chevrons reduces noise from both regions. This shows that the spatial zone of influence of each chevron is distinct, further supporting the notion that their effects might be independent

Figure 6b shows data for a Strouhal number of 9.1. This corresponds to the peak high frequency penalty from the far-field data. The spatial separation is not as clear at this Strouhal number, but the trend is still evident. Core chevrons alone make a lobe of additional noise at around two equivalent diameters. Fan chevrons alone make a region of increased noise at around three equivalent diameters. Applying both chevron nozzles increases the noise at both locations.

Acoustic results have demonstrated that core and fan chevrons act independently and that the noise reduction is additive. The suggestion that core and fan chevron nozzles modify noise independently is further supported by PIV results.14 Figure 7 shows the time-averaged axial velocities at several streamwise locations. It can be clearly seen that when core chevrons are applied, the velocity profiles become less steep in the region of the core shear layer, while in the cases where the fan chevrons are applied the fan shear layer mixes more rapidly. It can also be seen in


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Figure 8 that the Turbulent kinetic energy (TKE) increases significantly in the region just aft of each chevron nozzle when it is applied. It is less obvious, but certainly noticeable that the core chevrons reduce TKE downstream in the core/fan shear layer, and that fan chevrons reduce TKE downstream in the fan/ambient shear layer. The regions of influence of the two chevrons are spatially distinct. This further supports the notion that their acoustic effects would be independent as well.

B. Supersonic chevrons, Shock Cell noise When aircraft operate at supersonic cycle conditions with fixed convergent nozzles, they will generate

imperfectly expanded jets. With conic nozzles like those found on transport aircraft, the common condition at high cruise is an underexpanded jet. This creates repeated expansion fans and shocks as the flow adjusts itself to the ambient pressure. These conditions generate two additional noise producing mechanisms: Screech and broadband shock associated noise. These noise mechanisms do not usually affect the airport environment, however, at cruise the additional noise produced is annoying to passengers.

Broadband shock associated noise is created as turbulent structures pass through the quasi periodic series of shock cells of an imperfectly expanded jet. This interaction with successive shock cells results in constructive/destructive interference that creates preferred frequencies depending upon the observer’s directivity angle. A series of tests have been undertaken at UC to study the effects of chevrons on this noise source.15,16

Figure 9 shows the near-field and PIV results for a baseline conical nozzle and an eight-lobed chevron nozzle. The core NPR is 2.36 which with the convergent nozzle, yielded an under expanded jet, creating a series of quasi-periodic shock cells with a peak (fully expanded) Mach number of 1.18. The fan was run at a Mach number of 0.85. The near-field plots are for 6300Hz. This frequency is near the peak of the shock associated noise as determined by far-field measurement. It can be seen that the chevrons move the shock associated noise farther upstream, and reduce its intensity. The PIV results and the shadograph images reveal that the chevrons increase the number of shocks, and therefore reduce the strength of each individual shock.

C. Long Duct Chevron Mixer Another novel application of chevrons involves replacing the lobed mixer within mixed-flow exhaust systems

with a chevron nozzle. The resulting arrangement is known as a Long Duct Chevron Mixer (LDCM) and is shown in Figure 10. Tests at the UC-ATF show this configuration to yield an acoustic benefit.

Eight-lobed chevron nozzles with two different penetrations were compared to baseline conic nozzles at two different bypass ratios within a long duct. The hardware for the two different bypass ratios (five and eight) are of different lengths, so using the different bypass ratios also changes the length of the duct relative to the length of the nozzle within it. Application of chevron nozzles within the duct reduced the OASPL in all cases but one. In most cases the noise reduction was significant. (Figure 11). The long duct high penetration chevron nozzle configuration generally resulted in the greatest noise reduction.

Jet noise reduction was observed to be broadband, and most effective at lower frequencies. The directivity pattern of the overall sound pressure level was not significantly changed, however, a slight increase in higher frequencies was observed at aft angles due to the channeling of high frequency noise content in the direction of jet by the long duct nozzle.

The effectiveness of the long duct at reducing noise seemed to be dependent on the length of the duct since there were noticeable differences in the effectiveness of the LDCM between the two bypass ratios tested. In further investigations, particle imaging velocimetry testing could aid in determining the mechanism by which the long duct nozzle changes the flow field. It is hoped that this will yield a better understanding of the acoustic field and how to further reduce noise.17

D. Fluidic Injection Chevron nozzles have a significant disadvantage. They reduce thrust. Alternate technologies are sought which

will avoid this thrust penalty for the parts of the flight where acoustic treatment is unneeded. The acoustic benefit we seek is of great importance mostly while the aircraft is in the airport environment. If the thrust penalty were incurred only in the airport environment, the savings in fuel would be greatly beneficial. If we could develop a technology which can be turned off during cruise we would save greatly.

One technology which can be switched off in cruise is fluidic injection. Chevrons work by introducing streamwise vorticity into the shear layers behind the nozzle trailing edges. Fluidic injection aims to produce this vorticity by directly injecting small jets of air into the shear layer at angles. Fluidic injectors incur cost in that they will take bleed air from the engine, but once the aircraft has left the airport environment, the injectors will be turned off and the bleed will cease, thereby, removing the performance penalty.


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Initial tests were performed at the UC-ATF to examine the effect of such jets on single-stream core flow, and on the fan/ambient shear layer in dual flow.10 The hardware arrangement is shown in Figure 3a. Key parameters were identified in this single-flow arrangement, and a best arrangement was found to be tubes with 30° of pitch and 15° of yaw.10 Once this proof-of-concept testing was completed, a more involved test was planned and implemented at the GE Aviation Test Cell 41 Anechoic Free-Jet Test Facility (Figure 3b & c). This test was performed at larger scale and in three-stream flow. The third stream simulates forward movement of the vehicle.

The optimal configuration employs 26 core fluidic injection tubes and 28 fan injection tubes at 30° of pitch and 15° of yaw for both streams. The test demonstrated that jet noise reduction of 1.6 dB OASPL and 1.0 db EPNL can be achieved via steady fluidic injection at the nozzle exits with less than 1% of the total nozzle mass flow and at relatively low injection pressures. (Figure 12) The optimal fan momentum ratio was found to be 0.35%, which is smaller than the 1% reported by Callender10 at the UC-ATF in single-stream flow. This lower ratio may suggest that as the scale size increases, the required amount of fluidic injection does not scale with the mean jet diameter but perhaps scales with the boundary layer thickness. More work can be done to investigate the scaling effects as well as the mechanisms by which the fluidic injection alters the flow-field and acoustic near-field, thereby allowing for even greater optimization of the injection parameters.18

E. Numerical modeling Flows generating jet noise have a very large span of characteristic scales. The far-field contains very large

scales, while the scales responsible for acoustic sources are quite small. Handling such a large span of scales is a major difficulty from a computational point of view. In the field of Computational Aero-Acoustics (CAA) different approaches have been introduced to handle the different scales using problem-specific models and methods. The full, compressible flow governing equations can give full details of the analyzed unsteady flow field and the generated acoustics. However, solving the full equations of motion for compressible flows requires large computational time especially when far-field computations are performed. Additionally, there is no need for computing the flow itself in the far-field, where the contribution to acoustical sources is negligible. So, we use a decomposition technique in which the flow field is computed only where the acoustic sources are non-negligible, while only the acoustical field is computed in the far-field. In more exact terms, it is assumed that the velocity, the pressure, and the density can be decomposed into a flow related part and an acoustic fluctuation part. These parts are then solved separately using a flow solver (Navier-Stokes system of equations) and an acoustic solver (based on an inhomogeneous wave equation). The flow solver computes the flow field and provides the acoustic sources while the acoustic solver computes the acoustic wave propagation. The turbulence is modeled by using Large Eddy Simulation (LES) which is computationally less expensive than Direct Numerical Simulation (DNS), but it resolves the large energy-containing eddies. This captures the dynamics of turbulence and of the acoustic sources.

For the spatial discretization of the flow governing equations, a Cartesian staggered grid is used. The Cartesian grid is chosen for several reasons. The computational cost per cell is low. The convergence is fast. It is easy to implement high order discretization schemes. It has low storage requirements. The grid generation is fast.

For an efficient LES, attention must be paid to the grid quality. Very fine grid resolution is required in regions with large gradients, while in other parts of the domain a coarser grid resolution may be adequate. Here, this requirement is achieved by using local refinements that are easy to implement for Cartesian grids. The LES solver uses higher order schemes (i.e. more than two). Details of the numerical method are given in references 19-21.

The numerical approach was validated against experimental data in the case of a non-isothermal coaxial turbulent jet with and without chevron mixers on the core nozzle at a nominal shear condition. This case matches experimental measurements in the UC-ATF. The model tested was that of a medium by-pass turbofan engine. The core NPR was 1.85 and the shear velocity was 0.55 of the mass averaged mixed velocity. The core stream temperature was 240F and the fan stream temperature 90F.

Figure 13 (a) shows a comparison between the LES results and the experimental data for both cases in terms of normalized time-averaged axial velocity profiles along a radial line located at 2.5 equivalent diameters downstream from the fan exit nozzle plane. As can be seen, the computed data agrees well with the measurements at the specified location just downstream of the nozzle.

Downstream cross-section contour plots of time-averaged velocity magnitude and density computed by LES are presented for the coaxial nozzle case with chevrons on the core nozzle in Figure 13b & c. The acoustic predictions are shown in Figure 14. For both chevron and baseline nozzle the acoustic predictions are in very good agreement as far as trends are concerned. The only significant disparity between experiments and numerical data is at the aft-angles (directivity angle beyond 130°) far-field for the case when chevrons are used on the core nozzle.

Having validated the numerical method, we apply it to a new case, the case of an offset nozzle. In this case we have a heated core flow and an unheated fan flow surrounding it, but unlike the previously tested cases the two


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nozzles are not concentric. The inner nozzle is offset upward so that the exit from the outer fan nozzle is non-uniform in azimuth. (See reference 20) At the top of the nozzle there will be thin, high-speed stream. At he bottom there will be a thick low-speed stream. We would expect the radiated sound to vary azimuthally as well. We would expect the sound radiated upward, away from a ground-based observer to be increased, but the sound radiated downward toward the ground would be reduced. The far-field results below the nozzle are shown in Figure 15a. The predicted improvement in far-field noise is indeed significant, 4 dB improvement in OASPL at 90° (directly below the nozzle), diminishing to 1 dB at 150° (farther aft). Figure 15b shows the location of several lines along which near-field results are calculated. These results are shown in Figure 15c. The colors correspond to the colors of the arrows in subfigure b. In all cases the offset nozzle produces lower sound pressure levels than the conventional baseline configuration.

IV. Conclusion The University of Cincinnati has a large and productive aeroacoustics group, comprising both experimental and

computational aeroacoustics. The coaxial flow acoustic test rig provides a valuable and flexible facility to evaluate a variety of noise reducing concepts and prove them at small scale. Those which look promising are then selected by our industrial partners for testing at larger scale and higher fidelity.

Acknowledgments We wish to acknowledge General Electric Aircraft Engines of Cincinnati, Ohio, who provided funding for the

design and fabrication of the coaxial nozzle acoustic test rig, and funded much of the research presented in this paper.

References 1Peart, N.A., Smith, M.J.T., Magliozzi, B., Sternfeld, H., “Flyover-Noise Measurement and Prediction”. In “Aeroacoustics of

Flight Vehicles”, ed. H.H. Hubbard, Vol. 2, pp. 357-382, Acoustical Society of America, 1995. 2Lighthill, M.J., “On Sound Generated Aerodynamically. I. General Theory,” Proceedings of the Royal Society of London,

Series A, Volume 211, Issue 1107, pp 564-587, 1952. 3Lighthill, M.J., “On Sound Generated Aerodynamically. II. Turbulence as a source of sound,” Proceedings of the Royal

Society of London, Series A, Volume 222, Issue 1148, pp 1-32, 1954. 4Crow, S.C., Champagne, F.H., “Orderly Structure in Jet Turbulence,” Journal of Fluid Mechanics, Vol 48, Part 3, pp 547-

591, 1971. 5Brown, G.L., Roshko, A., “On Density Effects and Large Structure in Turbulent Mixing Layers,” Journal of Fluid

Mechanics, Vol 64, Part 4, pp 775-816, 1974. 6Tam, C.K.W., Golebiowski, M., & Seiner, J.M., “On the Two Components of Turbulent Mixing Noise from Supersonic

Jets,” AIAA 96-1716, 1996. 7Callender, W.B., Gutmark, E., Martens, S., “A Far-Field Acoustic Investigation of Chevron nozzle Mechanisms and

Trends,” AIAAJ Vol. 43, No. 1, pp. 87-95, 2005. 8Callender, W.B., Gutmark, E., Martens, S., “A Near Field Investigation of Chevron Nozzle Mechanisms,” AIAA 2003-

3210, 2003. 9Callender, W.B., Gutmark, E., Martens, S., “A PIV Flow Field Investigation of Chevron Nozzle Mechanisms,” AIAA 2004-

191, 2004. 10Callender, B. An investigation of innovative technologies for Reduction of Jet Noise in Medium and High Bypass Turbofan

Engines, PhD Dissertation, University of Cincinnati, 2004. 11Callender, B., Gutmark, E., DiMicco, R., “Design and Validation of a Coaxial Nozzle Acoustic Test Facility”, AIAA-2002-

0369, 2002. 12Rask, O. Gutmark, E., Martens, S. ”Acoustic Investigation of a High Bypass Ratio Separate Flow Exhaust System,” AIAA

2004-9, 2004. 13Rask, O., Gutmark, E., Martens, S., “A Near-Field Investigation of Primary and Secondary Chevron Nozzles,” AIAA

2007-437, 2006. 14Rask, O., Gutmark, E., Martens, S., “A PIV Flow Field Investigation of Medium Bypass Chevron Nozzles,” AIAA-2007-

436, 2007. 15Rask, O., Gutmark, E., Martens, S., “Broadband Shock Associated Noise Suppression by Chevrons,” AIAA-2006-9, 2006. 16Rask, O., Gutmark, E., Martens, S., “Shock Cell Modification due to Chevrons,’ AIAA-2007-831, 2007. 17Harrison, S.M., Gutmark, E., Martens, S., “A Far-Field Acoustic Investigation of Subsonic Jet Noise Reduction By Long

Duct Chevron Mixers (LDCM),” AIAA-2006-8, 2006. 18Harrison, S.M., Gutmark, E., Martens, S., “Jet Noise Reduction by Fluidic Injection On a Separate Flow Exhaust System,”

AIAA-2007-439, 2007.


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19Mihaescu, M., Gutmark, E., Szasz, R-Z., Fuchs, L., “Flow and Acoustics of a Coaxial Nozzle: a Sensitivity Study to the Inlet Boundary Conditions,” AIAA-2006-619, 2006.

20Mihaescu, M., Gutmark, E., Szasz, R-Z., Fuchs, L., “Computational Aeroacoustics of a Separate Flow Exhaust System with Eccentric Inner Nozzle,” AIAA-2007-13. 21Mihaescu, M., Gutmark, E., Fuchs, L., ” Computational Aeroacoustics of the Coaxial Flow Exhaust System of a Gas

Turbine Engine,” ASME-GT2007-28193, 2007.


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Figure 1 – UC-ATF coaxial flow acoustic rig. (a) cut-away view. (b) anechoic chamber lay-out. (c) complete rig. (d) nozzle close-up.

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Figure 4 – Spectral SPL results. a) complete spectra. b) Single Strouhal number SPL results for fan/corechevron combinations compared to baseline nozzle values.

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Figure 2 – Interchangeable nozzle hardware. (a) core nozzles. (b) fan nozzles. (a) (b)

Figure 3 – Fluidic injection hardware, (a) Initial UC-GDPL prototype test, single flow, small scale, (b) GE Cell 41 test, three flow, larger scale, (c) GE Cell 41 test.

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Figure 6 – Near-field mapping of subsonic, nominal-shear cycle condition. The cartoon indicates the presence or absence of chevrons in each case. (a) Strouhal = 0.45. (b) Strouhal = 9.1

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Figure 5 – Effect of Chevrons is additive between core and fan flows. The black trace is the arithmetic sum ofblue and green traces, illustrating the independent additive nature of the contributions from both sets ofchevrons.

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Figure 7 – PIV results for subsonic chevrons. (a) axial velocity at Deq=1.5. (b) axial velocity at Deq=3.0. (c) axial velocity at Deq=4.5. (d) axial velocity at Deq=6.0

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Figure 8 – Turbulent kinetic energy (TKE) results (a) both baseline nozzles. (b) with core chevrons. (c) withfan chevrons. (d) with chevrons on both core and fan.

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Figure 9 – Effect of chevrons on supersonic broadband noise. Primary nozzle is an underexpanded jet, with afully-expanded Mach number of 1.18. Secondary Mach = 0.85. Near-field plots are at 6300Hz which is the peak value for shock-cell noise. PIV and Shadograph images show that the chevrons increase the number ofshocks and reduce the strength of each shock.

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Figure 10 – Long Duct Chevron Mixer concept. (a) the Separate Flow (SF) configuration. (b) the Long Duct(LD) configuration.

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Figure 11 – Long Duct Chevron Mixer ∆OASPL. Legend is as follows 5-SF-8LP implies a bypass ratio of 5, Single flow, and the 8LP chevron nozzle. “SF” indicates separate flow, “LD” indicates the long duct configuration.

Figure 12 – Delta OASPL Comparison of Baseline and Optimal Fluidic Injection Cases for Core and Fan Injection Measured at a Directivity Angle of 150°


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Figure 13 – LES validation. (a) Computed radial profiles of time-averaged axial velocity as compared to experimental data for both cases. (b) Chevron nozzle: Time-averaged velocity field. (c) Chevron nozzle: Time-averaged density field.

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Figure 15 – Offset nozzle, acoustic predictions. (a) Far-field predictions. (b) Locations of near-field predictions. (c) Near-field predictions.

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Figure 14 – Baseline and Chevron, numerical acoustic predictions. (a) Far-field predictions compared to experiment. (b) Locations of near-field predictions. (c) Near-field predictions compared to experimental data.

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jet aircraft propulsion noise reduction

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