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NASA NTRS2000

Advanced Subsonic Technology (AST) Separate-Flow High-Bypass Ratio Nozzle Noise Reduction Program Test Report

NASA s model-scale nozzle noise tests show that it is possible to achieve a 3 EPNdB jet noise reduction with inwardfacing chevrons and flipper-tabs installed on the primary nozzle and fan nozzle chevrons. These chevrons and tabs are simple devices and are easy to be incorporated into existing short duct separate-flow nonmixed nozzle exhaust systems. However, these devices are expected to cause some small amount of thrust loss relative to the axisymmetric baseline nozzle system. Thus, it is important to have these devices further tested in a calibrated nozzle performance test facility to quantify the thrust performances of these devices. The choice of chevrons or tabs for jet noise suppression would most likely be based on the results of thrust loss performance tests to be conducted by Aero System Engineering (ASE) Inc. It is anticipated that the most promising concepts identified from this program will be validated in full scale engine tests at both Pratt & Whitney and Allied-Signal, under funding from NASA s Engine Validation of Noise Reduction Concepts (EVNRC) programs. This will bring the technology readiness level to the point where the jet noise suppression concepts could be incorporated with high confidence into either new or existing turbofan engines having short-duct, separate-flow nacelles.

Low, John K. C.↗

NASA NTRS2004

Experimental, Theoretical, and Computational Investigation of Separated Nozzle Flows

A detailed experimental, theoretical, and computational study of separated nozzle flows has been conducted. Experimental testing was performed at the NASA Langley 16-Foot Transonic Tunnel Complex. As part of a comprehensive static performance investigation, force, moment, and pressure measurements were made and schlieren flow visualization was obtained for a sub-scale, non-axisymmetric, two-dimensional, convergent- divergent nozzle. In addition, two-dimensional numerical simulations were run using the computational fluid dynamics code PAB3D with two-equation turbulence closure and algebraic Reynolds stress modeling. For reference, experimental and computational results were compared with theoretical predictions based on one-dimensional gas dynamics and an approximate integral momentum boundary layer method. Experimental results from this study indicate that off-design overexpanded nozzle flow was dominated by shock induced boundary layer separation, which was divided into two distinct flow regimes; three- dimensional separation with partial reattachment, and fully detached two-dimensional separation. The test nozzle was observed to go through a marked transition in passing from one regime to the other. In all cases, separation provided a significant increase in static thrust efficiency compared to the ideal prediction. Results indicate that with controlled separation, the entire overexpanded range of nozzle performance would be within 10% of the peak thrust efficiency. By offering savings in weight and complexity over a conventional mechanical exhaust system, this may allow a fixed geometry nozzle to cover an entire flight envelope. The computational simulation was in excellent agreement with experimental data over most of the test range, and did a good job of modeling internal flow and thrust performance. An exception occurred at low nozzle pressure ratios, where the two-dimensional computational model was inconsistent with the three-dimensional separation observed in the experiment. In general, the computation captured the physics of the shock boundary layer interaction and shock induced boundary layer separation in the nozzle, though there were some differences in shock structure compared to experiment. Though minor, these differences could be important for studies involving flow control or thrust vectoring of separated nozzles. Combined with other observations, this indicates that more detailed, three-dimensional computational modeling needs to be conducted to more realistically simulate shock-separated nozzle flows.

Hunter, Craig A.↗

NASA NTRS2005

Results of Aero/Acoustic Tests and Analytical Studies of a Two-Dimensional Eight-Lobe Mixer-Ejector Exhaust Nozzle at Takeoff Conditions

Hot flow aero-acoustic tests were conducted with Pratt & Whitney's High-Speed Civil Transport (HSCT) Mixer-Ejector Exhaust Nozzles by General Electric Aircraft Engines (GEAE) in the GEAE Anechoic Freejet Noise Facility (Cell 41) located in Evendale, Ohio. The tests evaluated the impact of various geometric and design parameters on the noise generated by a two-dimensional (2-D) shrouded, 8-lobed, mixer-ejector exhaust nozzle. The shrouded mixer-ejector provides noise suppression by mixing relatively low energy ambient air with the hot, high-speed primary exhaust jet. Additional attenuation was obtained by lining the shroud internal walls with acoustic panels, which absorb acoustic energy generated during the mixing process. Two mixer designs were investigated, the high mixing "vortical" and aligned flow "axial", along with variations in the shroud internal mixing area ratios and shroud length. The shrouds were tested as hardwall or lined with acoustic panels packed with a bulk absorber. A total of 21 model configurations at 1:11.47 scale were tested. The models were tested over a range of primary nozzle pressure ratios and primary exhaust temperatures representative of typical HSCT aero thermodynamic cycles. Static as well as flight simulated data were acquired during testing. A round convergent unshrouded nozzle was tested to provide an acoustic baseline for comparison to the test configurations. Comparisons were made to previous test results obtained with this hardware at NASA Glenn's 9- by 15-foot low-speed wind tunnel (LSWT). Laser velocimetry was used to investigate external as well as ejector internal velocity profiles for comparison to computational predictions. Ejector interior wall static pressure data were also obtained. A significant reduction in exhaust system noise was demonstrated with the 2-D shrouded nozzle designs.

Harrington, Douglas↗

NASA NTRS2001

Development of a Hybrid RANS/LES Method for Turbulent Mixing Layers

Significant research has been underway for several years in NASA Glenn Research Center's nozzle branch to develop advanced computational methods for simulating turbulent flows in exhaust nozzles. The primary efforts of this research have concentrated on improving our ability to calculate the turbulent mixing layers that dominate flows both in the exhaust systems of modern-day aircraft and in those of hypersonic vehicles under development. As part of these efforts, a hybrid numerical method was recently developed to simulate such turbulent mixing layers. The method developed here is intended for configurations in which a dominant structural feature provides an unsteady mechanism to drive the turbulent development in the mixing layer. Interest in Large Eddy Simulation (LES) methods have increased in recent years, but applying an LES method to calculate the wide range of turbulent scales from small eddies in the wall-bounded regions to large eddies in the mixing region is not yet possible with current computers. As a result, the hybrid method developed here uses a Reynolds-averaged Navier-Stokes (RANS) procedure to calculate wall-bounded regions entering a mixing section and uses a LES procedure to calculate the mixing-dominated regions. A numerical technique was developed to enable the use of the hybrid RANS-LES method on stretched, non-Cartesian grids. With this technique, closure for the RANS equations is obtained by using the Cebeci-Smith algebraic turbulence model in conjunction with the wall-function approach of Ota and Goldberg. The LES equations are closed using the Smagorinsky subgrid scale model. Although the function of the Cebeci-Smith model to replace all of the turbulent stresses is quite different from that of the Smagorinsky subgrid model, which only replaces the small subgrid turbulent stresses, both are eddy viscosity models and both are derived at least in part from mixing-length theory. The similar formulation of these two models enables the RANS and LES equations to be solved with a single solution scheme and computational grid. The hybrid RANS-LES method has been applied to a benchmark compressible mixing layer experiment in which two isolated supersonic streams, separated by a splitter plate, provide the flows to a constant-area mixing section. Although the configuration is largely two dimensional in nature, three-dimensional calculations were found to be necessary to enable disturbances to develop in three spatial directions and to transition to turbulence. The flow in the initial part of the mixing section consists of a periodic vortex shedding downstream of the splitter plate trailing edge. This organized vortex shedding then rapidly transitions to a turbulent structure, which is very similar to the flow development observed in the experiments. Although the qualitative nature of the large-scale turbulent development in the entire mixing section is captured well by the LES part of the current hybrid method, further efforts are planned to directly calculate a greater portion of the turbulence spectrum and to limit the subgrid scale modeling to only the very small scales. This will be accomplished by the use of higher accuracy solution schemes and more powerful computers, measured both in speed and memory capabilities.

Georgiadis, Nicholas J.↗

NASA NTRS2006

Phased-Array Study of Dual-Flow Jet Noise: Effect of Nozzles and Mixers

A 16-microphone linear phased-array installed parallel to the jet axis and a 32-microphone azimuthal phased-array installed in the nozzle exit plane have been applied to identify the noise source distributions of nozzle exhaust systems with various internal mixers (lobed and axisymmetric) and nozzles (three different lengths). Measurements of velocity were also obtained using cross-stream stereo particle image velocimetry (PIV). Among the three nozzle lengths tested, the medium length nozzle was the quietest for all mixers at high frequency on the highest speed flow condition. Large differences in source strength distributions between nozzles and mixers occurred at or near the nozzle exit for this flow condition. The beamforming analyses from the azimuthal array for the 12-lobed mixer on the highest flow condition showed that the core flow and the lobe area were strong noise sources for the long and short nozzles. The 12 noisy spots associated with the lobe locations of the 12-lobed mixer with the long nozzle were very well detected for the frequencies 5 KHz and higher. Meanwhile, maps of the source strength of the axisymmetric splitter show that the outer shear layer was the most important noise source at most flow conditions. In general, there was a good correlation between the high turbulence regions from the PIV tests and the high noise source regions from the phased-array measurements.

Soo Lee, Sang↗

NASA NTRS1978

Heat Resistant Paint

The racing car shown is one of many coated with an inorganic paint that protects "hot parts" of automotive vehicles. Developed and manufactured by Sperex Corporation, Gardena, California, the durable, heat-resistant paint is used on car and truck exhaust systems, firewalls, brake drums and engine manifolds. NASA technology contributed to development of the paint. Sperex was provided a technical support packa'ge detailing the research of Goddard Space Flight Center on long-life inorganic coatings. The information helped Sperex perfect its own formulations.

Source record↗

NASA NTRS2011

PIV Measurements of Chevrons on F400-Series Tactical Aircraft Nozzle Model

Reducing noise of tactical jet aircraft has taken on fresh urgency as core engine technologies allow higher specific-thrust engines and as society become more concerned for the health of its military workforce. Noise reduction on this application has lagged the commercial field as incentives for quieting military aircraft have not been as strong as in their civilian counterparts. And noise reduction strategies employed on civilian engines may not be directly applicable due to the differences in exhaust system architecture and mission. For instance, the noise reduction technology of chevrons, examined in this study, will need to be modified to take into account the special features of tactical aircraft nozzles. In practice, these nozzles have divergent slats that are tied to throttle position, and at take off the jet flow is highly overexpanded as the nozzle is optimized for cruise altitude rather than sea level. In simple oil flow visualization experiments conducted at the onset of the current test program flow barely stays attached at end of nozzle at takeoff conditions. This adds a new twist to the design of chevrons. Upon reaching the nozzle exit the flow shrinks inward radially, meaning that for a chevron to penetrate the flow it must extend much farther away from the baseline nozzle streamline. Another wrinkle is that with a variable divergence angle on the nozzle, the effective penetration will differ with throttle position and altitude. The final note of realism introduced in these experiments was to simulate the manner in which bypass flow is bled into the nozzle wall in real engines to cool the nozzle, which might cause very fat boundary layer at exit. These factors, along with several other issues specific to the application of chevrons to convergent-divergent nozzles have been explored with particle image velocimetry measurements and are presented in this paper.

Bridges, James↗

NASA NTRS2012

Aeroacoustics of Three-Stream Jets

Results from acoustic measurements of noise radiated from a heated, three-stream, co-annular exhaust system operated at subsonic conditions are presented. The experiments were conducted for a range of core, bypass, and tertiary stream temperatures and pressures. The nozzle system had a fan-to-core area ratio of 2.92 and a tertiary-to-core area ratio of 0.96. The impact of introducing a third stream on the radiated noise for third-stream velocities below that of the bypass stream was to reduce high frequency noise levels at broadside and peak jet-noise angles. Mid-frequency noise radiation at aft observation angles was impacted by the conditions of the third stream. The core velocity had the greatest impact on peak noise levels and the bypass-to-core mass flow ratio had a slight impact on levels in the peak jet-noise direction. The third-stream jet conditions had no impact on peak noise levels. Introduction of a third jet stream in the presence of a simulated forward-flight stream limits the impact of the third stream on radiated noise. For equivalent ideal thrust conditions, two-stream and three-stream jets can produce similar acoustic spectra although high-frequency noise levels tend to be lower for the three-stream jet.

Henderson, Brenda↗

NASA NTRS2012

Aeroacoustics of Three-Stream Jets

Henderson, Brenda S.↗

NASA NTRS2015

A Model for Jet-Surface Interaction Noise Using Physically Realizable Upstream Turbulence Conditions

This paper is a continuation of previous work in which a generalized Rapid Distortion Theory (RDT) formulation was used to model low-frequency trailing-edge noise. The research was motivated by proposed next-generation aircraft configurations where the exhaust system is tightly integrated with the airframe. Data from recent experiments at NASA on the interaction between high-Reynolds-number subsonic jet flows and an external flat plate showed that the power spectral density (PSD) of the far-field pressure underwent considerable amplification at low frequencies. For example, at the 900 observation angle, the low-frequency noise could be as much as 10dB greater than the jet noise itself. In this paper, we present predictions of the noise generated by the interaction of a rectangular jet with the trailing edge of a semi-infinite flat plate. The calculations are based on a formula for the acoustic spectrum of this noise source derived from an exact formal solution of the linearized Euler equations involving (in this case) one arbitrary convected scalar quantity and a Rayleigh equation Green's function. A low-frequency asymptotic approximation for the Green's function based on a two-dimensional mean flow is used in the calculations along with a physically realizable upstream turbulence spectrum, which includes a finite de-correlation region. Numerical predictions, based on three-dimensional RANS solutions for a range of subsonic acoustic Mach number jets and nozzle aspect ratios are compared with experimental data. Comparisons of the RANS results with flow data are also presented for selected cases. We find that a finite decorrelation region increases the low-frequency algebraic decay (the low frequency "rolloff") of the acoustic spectrum with angular frequency thereby producing much closer agreement with noise data for Strouhal numbers less than 0.1. Secondly, the large-aspectratio theory is able to predict the low-frequency amplification due to the jet-edge interaction reasonably well, even for moderate aspect ratio nozzles. We show also that the noise predictions for smaller aspect ratio jets can be fine-tuned using the appropriate RANS-based mean flow and turbulence properties.

Jet↗

NASA NTRS2015

Measurements and Predictions of the Noise from Three-Stream Jets

An experimental and numerical investigation of the noise produced by high-subsonic and supersonic three-stream jets was conducted. The exhaust system consisted of externally-mixed-convergent nozzles and an external plug. Bypass- and tertiary-to-core area ratios between 1.0 and 2.5, and 0.4 and 1.0, respectively, were studied. Axisymmetric and offset tertiary nozzles were investigated for heated and unheated conditions. For axisymmetric configurations, the addition of the third stream was found to reduce peak- and high-frequency acoustic levels in the peak-jet-noise direction, with greater reductions at the lower bypass-to-core area ratios. For the offset configurations, an offset duct was found to decrease acoustic levels on the thick side of the tertiary nozzle relative to those produced by the simulated two-stream jet with up to 8 dB mid-frequency noise reduction at large angles to the jet inlet axis. Noise reduction in the peak-jet-noise direction was greater for supersonic core speeds than for subsonic core speeds. The addition of a tertiary nozzle insert used to divert the third-stream jet to one side of the nozzle system provided no noise reduction. Noise predictions are presented for selected cases using a method based on an acoustic analogy with mean flow interaction effects accounted for using a Green's function, computed in terms of its coupled azimuthal modes for the offset cases, and a source model previously used for round and rectangular jets. Comparisons of the prediction results with data show that the noise model predicts the observed increase in low-frequency noise with the introduction of a third, axisymmetric stream, but not the high-frequency reduction. For an offset third stream, the model predicts the observed trend of decreased sound levels on the thick side of the jet compared with the thin side, but the predicted azimuthal variations are much less than those seen in the data. Also, the shift of the spectral peak to lower frequencies with increasing polar angle is over-predicted. For an offset third stream with a heated core, it is shown that including the enthalpy-flux source terms in the acoustic analogy model improves predictions compared with those obtained using only the momentum flux.

Three-Stream Jets↗

NASA NTRS2015

Measurements and Predictions of the Noise from Three-Stream Jets

An experimental and numerical investigation of the noise produced by high-subsonic and supersonic three-stream jets was conducted. The exhaust system consisted of externally-mixed-convergent nozzles and an external plug. Bypass- and tertiary- to-core area ratios between 1.0 and 2.5, and 0.4 and 1.0, respectively, were studied. Axisymmetric and offset tertiary nozzles were investigated for heated and unheated conditions. For axisymmetric configurations, the addition of the third stream was found to reduce peak- and high-frequency acoustic levels in the peak-jet-noise direction, with greater reductions at the lower bypass-to-core area ratios. For the offset configurations, an offset duct was found to decrease acoustic levels on the thick side of the tertiary nozzle relative to those produced by the simulated two-stream jet with up to 8 dB mid-frequency noise reduction at large angles to the jet inlet axis. Noise reduction in the peak-jet-noise direction was greater for supersonic core speeds than for subsonic core speeds. The addition of a tertiary nozzle insert used to divert the third-stream jet to one side of the nozzle system provided no noise reduction. Noise predictions are presented for selected cases using a method based on an acoustic analogy with mean flow interaction effects accounted for using a Green's function, computed in terms of its coupled azimuthal modes for the offset cases, and a source model previously used for round and rectangular jets. Comparisons of the prediction results with data show that the noise model predicts the observed increase in low-frequency noise with the introduction of a third, axisymmetric stream, but not the high-frequency reduction. For an offset third stream, the model predicts the observed trend of decreased sound levels on the thick side of the jet compared with the thin side, but the predicted azimuthal variations are much less than those seen in the data. Also, the shift of the spectral peak to lower frequencies with increasing polar angle is over-predicted. For an offset third stream with a heated core, it is shown that including the enthalpy-flux source terms in the acoustic analogy model improves predictions compared with those obtained using only the momentum- flux.

Three-Stream Jets↗

NASA NTRS2015

A Model for Jet-Surface Interaction Noise Using Physically Realizable Upstream Turbulence Conditions

Aircraft↗

NASA NTRS2016

A Model for Jet-Surface Interaction Noise Using Physically Realizable Upstream Turbulence Conditions

This paper is a continuation of previous work in which a generalized Rapid Distortion Theory (RDT) formulation was used to model low-frequency trailing-edge noise. The research was motivated by proposed next-generation aircraft configurations where the exhaust system is tightly integrated with the airframe. Data from recent experiments at NASA on the interaction between high-Reynolds-number subsonic jet flows and an external flat plate showed that the power spectral density (PSD) of the far-field pressure underwent considerable amplification at low frequencies. For example, at the 90deg observation angle, the low-frequency noise could be as much as 10 dB greater than the jet noise itself. In this paper, we present predictions of the noise generated by the interaction of a rectangular jet with the trailing edge of a semi-infinite flat plate. The calculations are based on a formula for the acoustic spectrum of this noise source derived from an exact formal solution of the linearized Euler equations involving (in this case) one arbitrary convected scalar quantity and a Rayleigh equation Green's function. A low-frequency asymptotic approximation for the Green's function based on a two-dimensional mean flow is used in the calculations along with a physically realizable upstream turbulence spectrum, which includes a finite decorrelation region. Numerical predictions of the sound field, based on three-dimensional RANS solutions to determine the mean flow, turbulent kinetic energy and turbulence length and time scales, for a range of subsonic acoustic Mach number jets and nozzle aspect ratios are compared with experimental data. Comparisons of the RANS results with flow data are also presented for selected cases. We find that a finite decorrelation region in the turbulence spectrum increases the low-frequency algebraic decay (the low frequency "roll-off") of the acoustic spectrum with angular frequency thereby producing much closer agreement with noise data for Strouhal numbers less than 0.1. Secondly, the large-aspect-ratio theory is able to predict the low-frequency amplification due to the jet-edge interaction reasonably well, even for moderate aspect ratio nozzles. We show also that the noise predictions for smaller aspect ratio jets can be fine-tuned using the appropriate RANS-based mean flow and turbulence properties.

Jet↗

NASA NTRS2016

Characterization of Three-Stream Jet Flow Fields

Flow-field measurements were conducted on single-, dual- and three-stream jets using two-component and stereo Particle Image Velocimetry (PIV). The flow-field measurements complimented previous acoustic measurements. The exhaust system consisted of externally-plugged, externally-mixed, convergent nozzles. The study used bypass-to-core area ratios equal to 1.0 and 2.5 and tertiary-to-core area ratios equal to 0.6 and 1.0. Axisymmetric and offset tertiary nozzles were investigated for heated and unheated high-subsonic conditions. Centerline velocity decay rates for the single-, dual- and three-stream axisymmetric jets compared well when axial distance was normalized by an equivalent diameter based on the nozzle system total exit area. The tertiary stream had a greater impact on the mean axial velocity for the small bypass-to-core area ratio nozzles than for large bypass-to-core area ratio nozzles. Normalized turbulence intensities were similar for the single-, dual-, and three-stream unheated jets due to the small difference (10%) in the core and bypass velocities for the dual-stream jets and the low tertiary velocity (50% of the core stream) for the three-stream jets. For heated jet conditions where the bypass velocity was 65% of the core velocity, additional regions of high turbulence intensity occurred near the plug tip which were not present for the unheated jets. Offsetting the tertiary stream moved the peak turbulence intensity levels upstream relative to those for all axisymmetric jets investigated.

Noise↗

NASA NTRS2016

Characterization of Three-Stream Jet Flow Fields

Flow-field measurements were conducted on single-, dual- and three-stream jets using two-component and stereo Particle Image Velocimetry (PIV). The flow-field measurements complimented previous acoustic measurements. The exhaust system consisted of externally-plugged, externally-mixed, convergent nozzles. The study used bypass-to-core area ratios equal to 1.0 and 2.5 and tertiary-to-core area ratios equal to 0.6 and 1.0. Axisymmetric and offset tertiary nozzles were investigated for heated and unheated high-subsonic conditions. Centerline velocity decay rates for the single-, dual- and three-stream axisymmetric jets compared well when axial distance was normalized by an equivalent diameter based on the nozzle system total exit area. The tertiary stream had a greater impact on the mean axial velocity for the small bypass-to-core area ratio nozzles than for large bypass-to-core area ratio nozzles. Normalized turbulence intensities were similar for the single-, dual-, and three-stream unheated jets due to the small difference (10 percent) in the core and bypass velocities for the dual-stream jets and the low tertiary velocity (50 percent of the core stream) for the three-stream jets. For heated jet conditions where the bypass velocity was 65 percent of the core velocity, additional regions of high turbulence intensity occurred near the plug tip which were not present for the unheated jets. Offsetting the tertiary stream moved the peak turbulence intensity levels upstream relative to those for all axisymmetric jets investigated.

Noise↗

NASA NTRS2019

Noise Measurements of a Low-Noise Top-Mounted Propulsion Installation for a Supersonic Airliner

A model-scale exhaust system was tested to validate low-noise concepts and noise prediction methods. The tests involved far-field acoustics, translating phased array, and particle image velocimetry; this report covers the far-field acoustic measurements. Data were acquired for a series of nozzles with different chevron designs, both uninstalled and installed on a representative aircraft planform. The impact of the various chevron treatments on the far-field noise was documented, along with the impact of the pylon and planform. For the baseline nozzle, installation produced a 2EPNdB reduction, as assumed in system studies. Chevrons were used to shift noise sources upstream to maximize the installation benefits and to reduce unshielded sources downstream. These resulted in reductions of 4-5EPNdB relative to the uninstalled baseline nozzle. Detailed analysis of spectral directivities behind the integrated EPNL metric gave insight into how well these concepts actually work. When correlated with particle image velocimetry measurements and phased array measurements, reported in companion papers, the explanation of acoustic benefits from top-mounted propulsion is clear as is the path toward optimization of the concept.

Supersonics↗

NASA NTRS2019

PIV Measurements of a Low-noise Top-mounted Propulsion Installation for a Supersonic Airliner

A model-scale exhaust system was tested to validate low-noise propulsion concepts and noise prediction methods. The tests involved far-field acoustics, phased array, and particle image velocimetry (PIV). This paper covers the particle image velocimetry portion. Data was acquired at NASA Glenn's Aero-Acoustic Propulsion Lab for a series of nozzles with different chevron designs, both uninstalled and installed on a representative aircraft planform. The impact of the various chevron treatments on the turbulent velocity field was documented, along with the impact of the pylon and planform. When correlated with far-field acoustic measurements and phased array measurements, reported in companion papers, the explanation of acoustic benefits from top-mounted propulsion is clear as is the path toward optimization of the concept.

Propulsion Airframe Integration↗