Brenda S. Henderson, James E. Bridges, and Mark P. Wernet

Glenn Research Center, Cleveland, Ohio

Jet Noise Reduction Potential From Emerging

Variable Cycle Technologies

NASA/TM—2012-217744

November 2012
NASA STI Program . . . in Profi le

Since its founding, NASA has been dedicated to the 

advancement of aeronautics and space science. The 

NASA Scientifi c and Technical Information (STI) 

program plays a key part in helping NASA maintain 

this important role.

The NASA STI Program operates under the auspices 

of the Agency Chief Information Offi cer. It collects, 

organizes, provides for archiving, and disseminates 

NASA’s STI. The NASA STI program provides access 

to the NASA Aeronautics and Space Database and 

its public interface, the NASA Technical Reports 

Server, thus providing one of the largest collections 

of aeronautical and space science STI in the world. 

Results are published in both non-NASA channels 

and by NASA in the NASA STI Report Series, which 

includes the following report types:

• TECHNICAL PUBLICATION. Reports of 

completed research or a major signifi cant phase 

of research that present the results of NASA 

programs and include extensive data or theoretical 

analysis. Includes compilations of signifi cant 

scientifi c and technical data and information 

deemed to be of continuing reference value. 

NASA counterpart of peer-reviewed formal 

professional papers but has less stringent 

limitations on manuscript length and extent of 

graphic presentations.

• TECHNICAL MEMORANDUM. Scientifi c 

and technical fi ndings that are preliminary or 

of specialized interest, e.g., quick release 

reports, working papers, and bibliographies that 

contain minimal annotation. Does not contain 

extensive analysis.

• CONTRACTOR REPORT. Scientifi c and 

technical fi ndings by NASA-sponsored 

contractors and grantees.

• CONFERENCE PUBLICATION. Collected 

papers from scientifi c and technical 

conferences, symposia, seminars, or other 

meetings sponsored or cosponsored by NASA.

• SPECIAL PUBLICATION. Scientifi c, 

technical, or historical information from 

NASA programs, projects, and missions, often 

concerned with subjects having substantial 

public interest.

• TECHNICAL TRANSLATION. Englishlanguage translations of foreign scientifi c and 

technical material pertinent to NASA’s mission.

Specialized services also include creating custom 

thesauri, building customized databases, organizing 

and publishing research results.

For more information about the NASA STI 

program, see the following:

• Access the NASA STI program home page at 

http://www.sti.nasa.gov

• E-mail your question to help@sti.nasa.gov

• Fax your question to the NASA STI 

Information Desk at 443–757–5803

• Phone the NASA STI Information Desk at

 443–757–5802

• Write to:

 STI Information Desk

 NASA Center for AeroSpace Information

 7115 Standard Drive

 Hanover, MD 21076–1320
Brenda S. Henderson, James E. Bridges, and Mark P. Wernet

Glenn Research Center, Cleveland, Ohio

Jet Noise Reduction Potential From Emerging

Variable Cycle Technologies

NASA/TM—2012-217744

November 2012

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

48th Joint Propulsion Conference and Exhibit

cosponsored by the AIAA, ASME, SAE, and ASE

Atlanta, Georgia, July 30–August 1, 2012
Acknowledgments

The HVC hardware used in the experiments was designed and manufactured by Rolls-Royce Liberty Works under NASA’s 

Supersonics Project NRA NNL08AA29C. The N+2 HVC hardware was designed by Lockheed Martin and Rolls-Royce Liberty 

Works under Supersonics Project contract NNC10CA02C. The GEGR hardware was designed by Lockheed Martin and General 

Electric Global Research under Supersonics Project contract NNC10CA02C. This work was supported by the Supersonics Project 

in the Fundamental Aeronautics Program.

Available from

NASA Center for Aerospace Information

7115 Standard Drive

Hanover, MD 21076–1320

National Technical Information Service

5301 Shawnee Road

Alexandria, VA 22312

Available electronically at http://www.sti.nasa.gov

Trade names and trademarks are used in this report for identifi cation 

only. Their usage does not constitute an offi cial endorsement, 

either expressed or implied, by the National Aeronautics and 

Space Administration.

This work was sponsored by the Fundamental Aeronautics Program 

at the NASA Glenn Research Center.

Level of Review: This material has been technically reviewed by technical management. 
NASA/TM—2012-217744 1

Jet Noise Reduction Potential From Emerging

Variable Cycle Technologies

Brenda S. Henderson, James E. Bridges, and Mark P. Wernet

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Abstract

Acoustic and flow-field experiments were conducted on exhaust concepts for the next generation 

supersonic, commercial aircraft. The concepts were developed by Lockheed Martin (LM), Rolls-Royce 

Liberty Works (RRLW), and General Electric Global Research (GEGR) as part of an N+2 (next 

generation forward) aircraft system study initiated by the Supersonics Project in NASA’s Fundamental 

Aeronautics Program. The experiments were conducted in the Aero-Acoustic Propulsion Laboratory at 

the NASA Glenn Research Center. The exhaust concepts utilized ejectors, inverted velocity profiles, and 

fluidic shields. One of the ejector concepts was found to produce stagnant flow within the ejector and the 

other ejector concept produced discrete-frequency tones that degraded the acoustic performance of the 

model. The concept incorporating an inverted velocity profile and fluid shield produced overall-soundpressure-level reductions of 6 dB relative to a single stream nozzle at the peak jet noise angle for some 

nozzle pressure ratios. Flow separations in the nozzle degraded the acoustic performance of the inverted 

velocity profile model at low nozzle pressure ratios.

I. Introduction

The Supersonic Project of the NASA Fundamental Aeronautics Program is developing technologies 

for future, supersonic, commercial aircraft. The Project has established a set of sonic boom, airport noise, 

and emission goals for an N+2 aircraft (next generation forward) that could enter service between 2018 

and 2020. The N+2 aircraft are expected to achieve cruise speeds between Mach 1.6 and 1.8, have 

payloads between 35 and 70 passengers, and achieve airport noise levels 10 to 20 EPNdB (cumulative) 

below FAA Stage 3 requirements. N+2 systems studies conducted by Lockheed Martin (LM), RollsRoyce Liberty Works (RRLW), and General Electric Global Research (GEGR) indicated that a successful 

airplane design could be achieved if certain technologies performed as well as projected. The projected 

technologies to be validated used three-stream engine architectures and corresponding exhaust nozzles 

that could simultaneously achieve the low noise operation around airports and the high specific thrust 

required for supersonic cruise. To prove whether the proposed engine cycle and exhaust-nozzle concepts 

could meet the noise goal, model-scale versions of the nozzles were designed by RRLW and GEGR and 

tested at the NASA Glenn Research Center. The results of the N+2 exhaust-concept experiments are 

reported here.

The exhaust concepts proposed by RRLW incorporated ejectors (cylindrically shaped shrouds that 

surround the jet plume and often extend downstream of the nozzle exit for several nozzle diameters). 

Research on ejectors for jet-noise suppression began in the mid 1950s with investigations focusing on 

noise reduction potential for ejectors used with single-stream jets. Some of the first experiments identified 

the need of combining mixing nozzles with ejectors as ejectors were found to provide little or no 

suppression when combined with standard round nozzles (Refs. 1 to 3). Since these early investigations, 

most, if not all, ejector-exhaust designs incorporate mixing nozzles. These early investigations also 

identified the potential for production of discrete tones by ejectors. Initial parametric studies that would 

ultimately provide critical design information included engine and flight tests that focused on the 

importance of ejector diameter, length, and axial location (Refs. 4 and 5). The engine tests showed that up 

to 8 dB reduction could be achieved when ejectors were used with mixing nozzles and, in most cases, the 
NASA/TM—2012-217744 2

ejectors provided additional noise suppression over that achieved by the mixing nozzles alone as long as 

the ejector had sufficient length (typically greater than two nozzle diameters). Results from early flight

tests indicated that the mixer-ejector provided little or no in-flight perceived noise benefit although flight 

studies conducted later (Refs. 6 and 7) showed that perceived noise level reductions could be achieved 

with a combination of mixer and acoustically-lined ejector. These early studies focused on using the 

mixer-ejector to reduce jet noise through the reduction of jet velocity. It was realized that complete 

mixing of the primary jet with entrained ejector flow needed to occur before the exit of the ejector to 

adequately reduce the jet velocity and the ejector also needed to provide sufficient length to shield any 

high frequency noise introduced by the enhanced mixing process. A better physical understanding of the 

mixer-ejector flow was developed somewhat later when the ideal, maximum, ejector-pumping 

performance (mass flow entrainment) was shown (through an idealized application of the mass, 

momentum, and energy equations to the ejector control volume) to be related to the ejector area and 

occurred when the flows are fully mixed at the ejector exit (Refs. 8 to 10). Ejector pumping was shown to 

increase with increasing ejector area although, in real flows, an optimum area exists and increased area 

beyond this point produces diminished ejector pumping due to losses in the secondary duct which are not 

captured in the idealized ejector analysis. The idealized analysis does not address the mixing process 

within the ejector that is a function of ejector length and is the result of turbulent mixing between the

primary jet and ejector flow. When the turbulent shear layer formed by the primary jet and ejector flow 

extends to the ejector-shroud wall, a strong streamwise flow is established across the entire ejector cross 

section and significant suction is created at the ejector inlet. Enhanced mixing (and therefore shorter 

ejectors) can be achieved through mixers that streamwise vortices may allow for shortened ejector 

lengths. While the single-stream studies provide critical guidance for ejectors designed for future engine 

architectures, the flow structure of multi-stream jets incorporating mixing nozzles is far more complicated 

than the single stream counterpart. Mixers in multi-stream jets are typically used to mix the high-speed 

core flow with the fan flow and may not provide adequate mixing between the mixed core and fan streams 

with the entrained ejector flow to achieve adequate pumping of the ejector. The implementation of an 

ejector in the RRLW model provides the opportunity to evaluate ejectors in a multi-stream application.

The GEGR concept combined two noise-reduction concepts that had been explored previously: 

inverted velocity profiles (IVP) and fluid shields (FS). The inverted velocity profile concept takes the cold 

flow typically produced by the fan stage of the engine and redirects it along the centerline of the exhaust 

plume while diverting the hot stream to an annulus around the cold stream. Several reports (Refs. 11 to 

15) from the early 1980’s documented these impacts. This was considered beneficial for noise for several 

reasons. In previous incarnations of this concept the bypass ratio was relatively low and the cold bypass 

stream had a Mach number greater than 1, thus containing shocks from the underexpanded nozzle. 

Sheathing this shock-containing stream in the hot flow changed the broadband shock-associated noise 

produced. Additionally, by spreading the high speed flow over a thin annulus it mixed out sooner than 

when it was confined to the center of the jet, providing a lower mixed jet velocity at the peak jet mixing 

noise source region downstream. This effect was accentuated as the bypass ratio was increased. The price 

paid for enhanced mixing is increased noise from the initial shear that now has a higher velocity gradient. 

The fluid shield concept aims to reduce the high frequency penalty of the IVP by decreasing the higher 

velocity gradient over the first portion of the jet plume.

The results of the acoustic and flow-field studies for the RRLW and GEGR models are presented. The 

facilities, hardware, and measurements are described in the following section. Far-field acoustics were 

used to investigate the acoustic performance of the models and Particle Image Velocimetry measurements 

were used to investigate the flow fields. Results of the experiments are presented in Section III and a 

discussion of the concept development process is contained in Section IV.
NASA/TM—2012-217744 3

II. Experimental Approach

The experiments were conducted in the Aero-Acoustic Propulsion Laboratory (AAPL) at the NASA 

Glenn Research Center shown in Figure 1. The AAPL is a 65 ft radius geodesic dome treated with 

acoustic wedges. The AAPL contains the Nozzle Acoustic Test Rig (NATR), which produces a 53 in.

diameter simulated forward-flight stream reaching Mach numbers of 0.35. The High Flow Jet Exit Rig 

(HFJER), a dual-stream jet engine simulator capable of replicating most commercial turbo-fan engine 

temperatures and pressures (see Ref. 16), is centered in the simulated flight stream of the NATR. A thirdstream capable of achieving mass-flow rates between 0.5 to 6 lbm/sec has been added to the HFJER. The 

third-stream flow temperatures are equal to that of the bypass stream as the former stream branches off 

the latter stream after the bypass heater.

Acoustic measurements were made with the far-field array shown in Figure 1. The array contains

24 microphones located on a 45 ft constant radius arc covering polar angles between 45° and 160°, where 

angles greater than 90° are in the downstream direction relative to the third-stream nozzle exit. All data are 

corrected for atmospheric absorption (Ref. 17) and wind tunnel shear layer effects (Ref. 18) and are 

presented on a one-foot lossless (no atmospheric absorption) arc. Data are acquired using 1/4 in. Bruel and 

Kjaer microphones (without protection grids) pointed directly at the nozzle exit. Microphone sensitivity and 

frequency response have been applied to all measurements. Narrowband data were acquired in 12 Hz bands 

and results are presented as power-spectral density (PSD) unless otherwise stated.

Two types of particle image velocimetry (PIV) measurements were performed: (1) two-component 

streamwise experiments and (2) cross-stream stereo PIV experiments. For all PIV studies, Redlake 

ES11000 cameras with 2.6K x 4K CCD arrays were used to record images. For the streamwise 

experiments, two cameras, each with a 380 by 325 mm field of view, were used in a side-by-side 

configuration with a 28 mm overlap between camera images. The laser light sheet (created with a 400 mJ 

per pulse, 532 nm, dual-head Nd-YAG laser) illuminated the jet centerline. In the cross-stream 

experiments, the laser light sheet illuminated cross-stream planes in the jet. The field of view for the 

reconstructed velocity field was 400 by 360 mm. For all PIV experiments, the core and fan streams were 

seeded with a pH stabilized dispersion of aluminum oxide (0.7 m particles) in ethanol. The third stream 

was seeded with dry alumina particles using a fluidized bed seeder. The free jet was seeded with a Rosco 

Delta 3000 foggers that generates 0.5 to 4.5 m particles.

For the stereo PIV experiments, the 2-D vector maps from the left and right cameras were generated 

using a multi-pass processing approach. The processing strategy used an initial pass with 64 by 64 pixel 

subregions on 32 pixel centers followed by 6 passes (using simulated annealing) at 32 by 32 pixel 

subregions on 16 pixel centers, followed by two final passes using subregion distortion processing. 

Symmetric Phase Only Filtering was used on the cross-stream data sets to reduce the effect of flare light 

from the model.

All PIV surveys used 400 image pairs per axial station, yielding 400 velocity vector maps at each 

axial station. The velocity vector maps were then ensemble averaged to compute the first and second

order statistics. Hard velocity cutoff limits and Chauvenet’s criteria were used to remove outliers.

The RRLW models are shown in Figure 2. The model in Figure 2(a), the Highly Variable Cycle 

(HVC) exhaust model, was used in experiments preceding the N+2 system studies and served as the basis 

for the redesigned model (see Fig. 2(b)) used in the N+2 studies. The HVC model had a core lobed-mixer 

nozzle, an elliptic fan nozzle, and an actuated ejector. Sidewalls were used to attach the ejector doors to 

the outer shell of the fan nozzle. The N+2 HVC model shown in Figure 2(b) used the HVC core lobed 

mixer, a circular fan nozzle, a third-stream nozzle that pumped an ejector, and ejector flaps that diverted 

flow from the fan and third streams to the core stream. The contours of the ejector and sidewalls of the 

N+2 HVC model were significantly different from those used in the HVC model. The three-stream nozzle 

system in Figure 2(c) was the baseline-nozzle system for the N+2 HVC investigations.
NASA/TM—2012-217744 4

The cycle points used in the HVC experiments are shown in Table 1. The nozzle pressure ratio, NPR, 

is the ratio of the stagnation pressure of the jet to the ambient pressure. The nozzle temperature ratio, 

NTR, is the ratio of the stagnation temperature of the jet to the ambient temperature. Subscripts c and f 

indicate core and fan stream conditions, respectively. As shown in the Table, all data were acquired at 

subsonic exhaust speeds on the core and fan streams and heated conditions on the core stream. The data 

were acquired at free jet Mach numbers (Mfj) equal to 0 (static conditions) and 0.3.

The cycle points used in the N+2 HVC experiments were similar to those shown in Table 1 with the 

third stream operated at nozzle pressure ratios slightly lower than those of the fan stream. For all 

experiments, the exhaust speeds of the core, fan, and tertiary stream were subsonic.

A cross-sectional view of the GEGR model is shown in Figure 3. The core stream of the HFJER was 

internally routed within the model to the annular “Hot Flow” stream shown in Figure 3. The fan stream 

was routed to the inner flow stream along the plug. The fluid shield was on the measurement side of the 

jet was supplied from the HFJER tertiary stream. A conventional aspect of the nozzle was the changing of 

the divergent part of the nozzle. Plans for the full-scale engine had variable A9/A8 to optimize cruise 

performance. However, in the first design iteration two nozzle contours were designed, both of which 

were highly overexpanded at low-altitude operating points. Based on preliminary RANS CFD, it was felt 

that the acoustic effect of the overexpanded flow would be minimal. This was determined to be a mistake 

in validation tests. An additional difference from previous work was the mixing of streams to produce the 

“hot” outer stream. The effective bypass ratio of cold flow to hot flow was therefore less than 1, pushing 

the concept outside the range of previous experiments. Additionally, the nozzle features a significant 

plug, required to minimize the effect of the nozzle boat tail angle. The reference (baseline) configuration 

was a single-stream nozzle with the same fully-mixed flow conditions as those of the IVP configuration.

III. Results

A. RRLW HVC and N+2 HVC Model

Acoustic results for the HVC model (see Fig. 2(a)) operated at representative takeoff conditions and Mfj

= 0.0 are shown in Figure 4. The angles indicated in the legend are the ejector door angles relative to the 

closed position (no ejector opening) at 0°. As the ejector door angle increases, the ejector inlet area 

increases. Observation angles greater than 90° are in the downstream direction relative to the ejector-door 

trailing edge. The baseline data were acquired for an equivalent-area-nozzle system with a round fan nozzle 

and no ejector doors. Tones are present in the acoustic data acquired at the 5° door position for all 

observation angles. The baseline nozzle produced the lowest acoustic levels for observation angles less than, 

or equal to, 90°. In the peak jet noise direction (see Fig. 4(b)), acoustic levels for the 10° and 20° ejector 

door configurations are lower than those of the baseline-nozzle system for frequencies below 800 Hz.

Acoustic results obtained for the same jet conditions as those in Figure 4 and Mfj = 0.3 are shown in 

Figure 5. The tones observed in the spectra acquired at Mfj = 0.0 are not present in the spectra at Mfj = 0.3.

No acoustic benefit was achieved with the ejector relative to the baseline-nozzle system except at very 

low frequencies (below 400 Hz) in the peak jet noise direction.

The results from cross-stream stereo PIV experiments obtained at the same conditions as those for

Figure 4 are shown in Figure 6. The orientation of the model is shown in the upper left corner of the 

Figure. The location of the ejector-door trailing edges are indicated with solid black lines in the contour 

plot in the top row of Figure 6(a). The solid black arrows on the mean velocity contours indicate the 

velocity vector direction in that region of the flow. The mean velocity plots show that the corners of the 

ejector sidewalls distort the flow and create vorticity (not shown in the Figure) that elongate the jet plume 

in the direction of the sidewalls. Additionally, the fluid within the ejector appears to be stagnant. The 

vorticity in the core region of the flow generated by the core-lobed mixer is distorted by the elliptic fan 

nozzle and ejector. The distortion of the core flow inhibits the mixing of the core-lobed nozzle. The 

turbulent kinetic energy contours show that the elongation of the flow in the direction of the sidewalls 

increases TKE in the region of the shear layer directly downstream of the ejector doors. The inadequate 
NASA/TM—2012-217744 5

acoustic performance of the HVC model is likely associated with the lack of ejector flow and the 

increases in TKE levels associated with the impact of the ejector sidewalls on the evolving jet plume.

The results of PIV experiments at the same jet conditions as those indicated in Figure 6 with Mfj = 0.2 

are shown in Figure 7. The trends in the data at Mfj = 0.2 are similar to those observed at Mfj = 0.0. The 

introduction of the simulated forward flight stream did not induce flow in the ejector and the flow 

distortions associated with the ejector sidewalls and elliptical fan nozzle were similar to those observed at 

Mfj = 0.0.

Acoustic results for the N+2 HVC model operated at core and fan stream conditions similar to those 

in Figure 5 are shown in Figure 8. Multiple discrete tones were produced by the model at all observation 

angles. Measurements at other operating conditions and for other ejector and ejector flap angles produced 

results similar to those shown in Figure 8. Cavities between the ejector sidewalls and ejector doors as well 

as the ejector-flap hinges were covered and the measurements were repeated. These model modifications 

had little impact on the production of tones. Closing the ejector flap reduced or eliminated the tones but 

resulted in elevated broadband noise most likely due to flow separation in the core stream resulting from 

the divergence angle of the ejector flap. Subsequent experiments were conducted with the ejector flap 

deflected (resulting in a convergent ejector flap contour), the ejector flap opening covered, and the surface 

contour upstream of the ejector-flap entrance smoothed. The spectra in Figure 9, obtained after the 

ejector-flap modifications, show that the modifications eliminated the discrete tones indicating that the 

tones were the result of unsteady flow separations upstream of the ejector flap and within the ejector flap 

opening. It was not possible to eliminate discrete tones for all operating conditions and ejector flap angles.

Results from the N+2 HVC stereo PIV experiments are shown in Figure 10. The mean streamwise 

velocity contours indicate that ejector flow has been established for the N+2 HVC model. However, the 

modified ejector sidewalls did not eliminate the impact of the sidewalls on the downstream evolution of 

the flow which resulted in higher than expected TKE levels in the shear layer downstream of the region 

where the ejector doors connected to the ejector sidewalls. Additionally, the core-lobed mixer appeared to 

have little impact on the jet mixing.

B. GEGR Model

Acoustic results for the GEGR model are shown in Figure 11 and compared with those obtained 

reference nozzle (“REF”) for representative temperature cycles (“low”, “mid”, “hi”). For a constant 

temperature, the ideal thrust was increased by increasing NPR. The results show that the radiated noise 

levels associated with the inverted velocity profile drop dramatically above a given NPR (ideal thrust) and 

can be lower than the corresponding reference nozzle for the highest NPR investigated. However, for 

lower thrust points, the inverted velocity profile jet produced noise levels higher than those of the 

reference nozzle. PIV results indicated that increased noise levels (compared to the reference nozzle) at 

low NPR were probably generated by flow separation near the nozzle trailing edge. Figure 12 shows an 

instantaneous velocity vector map from the PIV testing superimposed on the nozzle. At the lip between 

the thin inner cold flow and larger hot outer there is a large low velocity region, the result of separation in 

the hot stream in the divergent portion of the nozzle. When statistics are computed from the velocity maps 

(see Fig. 13), the mean velocity shows a near-zero mean velocity in this region of the flow. There is also a 

significant amount of turbulence generated here, with peak levels greater than that found at the peakmixing region downstream. Although such separations and turbulence were found for a wide range of 

flow conditions, all of which were overexpanded for the divergent nozzle, there were significant 

differences in the details of the separation. Notably, initial RANS CFD solutions used to design the model 

flow lines did not pick up the strong flow separations resulting from the overexpanded flow. However, 

Large Eddy Simulations (LES) conducted after the test (see Fig. 3) clearly predicted the separation.

Acoustic results obtained at high enough nozzle-pressure ratios to suppress separation at the divergent 

nozzle exit (higher than those used in Fig. 3) are shown in Figures 14 and 15. The overall sound pressure 

levels (OASPL) for the IVP nozzle were 2 dB below those of the reference nozzle at polar angles 

broadside to the jet and 6dB below those of the reference nozzle at the peak radiation angle (see Fig. 14). 
NASA/TM—2012-217744 6

The sound spectra in Figure 15 show a 3 to 5 dB reduction in peak levels and a slight high-frequency 

increase for the IVP nozzle relative to the reference nozzle.

The fluid shield flow produced a further reduction of ~1 dB, but did not provide more reduction at the 

high frequencies where the IVP increased acoustic levels relative to levels produced by the reference 

nozzle. Most of this reduction was attributed to the reduction in fully mixed jet velocity, all cases being 

evaluated on a constant thrust basis.

IV. Discussion and Conclusions

The flow lines for both HVC models were based on results from RANS solutions. Due to the screening 

nature of the effort, many of the three-dimensional characteristics of the model (such as the lobed mixer 

geometry and ejector-sidewall contours) were omitted from the computation. The screening RANS solutions 

did not capture the flow separations and unwanted vorticity associated with realistic components in the 

design. While high fidelity RANS calculations performed on the HVC model following testing adequately 

captured the stagnant flow in the ejector and the general impact of the ejector sidewalls on the evolving jet 

plume, the computation and gridding time were significant and computational solutions of this type would 

be difficult to implement in an evolving design. Additionally, time-accurate solutions may be required to 

capture the unsteady separations associated with N+2 HVC model.

The test results for the GEGR model showed the potential for noise reduction from IVP and FS 

concepts. However, the test results also indicated the need for accurate computational flow solutions early 

in the design process. The RANS solutions did not capture the nozzle flow separations that resulted in 

poor acoustic performance at low NPR. Additionally, LES solutions were not available until after the 

model was designed and tested possibly due to the required computation time.

One major result of the studies presented here is the need for reliable and efficient computational 

tools for future low-noise exhaust concepts that involve complex geometries. Critical flow separations 

were not captured in screening RANS calculations and these flow separations resulted in reduced acoustic 

performance of both exhaust concepts. As flow streams are added to the engine, the complexity of the 

flow lines increases which increases the possibility of flow separation. In addition to reliable and efficient 

computations, there is a need to incorporate all relevant flow surfaces in the computational domain.

References

1. North, W., and Coles, W. “Effect of exhaust-nozzle ejectors on turbo jet noise generation,” NACA 

TN 3573, 1955.

2. Brausch, J.F. “Flight velocity influence on jet noise of conical ejector, annular plug and segmented 

suppressor nozzles,” NASA CR 120961, 1972.

3. Ciepluch, C., North, W., Coles, W., and Antl, R. “Acoustic, thrust, and drag characteristics of 

several full-scale noise suppressors for turbojet engines,” NACA TN 4261, 1958.

4. Coles, W., Mihaloew, J., and Callaghan, E. “Turbojet engine noise reduction with mixing nozzleejector combinations,” NACA TN 4317, 1958.

5. Coles, W., Mihaloew, J., and Swann, W. “Ground and in-flight acoustic and performance 

characteristics of jet-aircraft exhaust noise suppressors,” NASA TN D-874, 1961.

6. FitzSimmons, R.D, McKinnon, R.A., Johnson, E.S., and Brooks, J.R. “Flight and wind tunnel test 

results of mechanical jet noise suppressor nozzle,” AIAA-80-0165, 1980.

7. FitzSimmons, R.D., McKinnon, R.A., Johnson, E.S., and Brooks, J.R. “Flight wind-tunnel test 

results of a mechanical jet noise suppressor nozzle,” J. Aircraft, vol. 18, no. 9, 725 – 730, 1981.

8. Tillman, T.G., Paterson, R.W., and Presz, W.M. “Supersonic nozzle mixer ejector,” AIAA-89-2925, 

1989.

9. Presz, W., Reynolds, G. and Hunter, C. “Thrust augmentation with mixer/ejector systems,” AIAA 

2002-0230, 2002.
NASA/TM—2012-217744 7

10. Lord, W.K., Jones, C.W., Stern, A.M., Head, V.L., and Krejsa, E.A. “Mixer ejector nozzle for jet 

noise suppression,” AIAA 90-1909, 1990.

11. Goodykoontz, J.H. and Stone, J.R., “Experimental Study of Coaxial Nozzle Exhaust Noise,” NASA 

TM-79090 (1979).

12. Goodykoontz, J.H., “Inverted Velocity Profile Semi-Annular Nozzle Jet Exhaust Noise 

Experiments,” NASA (1985). 

13. Goodykoontz, J.H. & von Glahn, U., “Noise Suppression Due to Annulus Shaping of an InvertedVelocity-Profile Coaxial Nozzle,” NASA TM 81460 (1981).

14. Von Glahn, U.H. & Goodykoontz, J.H., “Plume Characteristi9cs of Single-Stream and Dual-Flow 

Conventional and Inverted-Profile Nozzles at Equal Thrust,” NASA TM 87323 (AIAA paper 86-

1809). 1986.

15. Tanna, H.K., Tester, B.J., and Lau, J.C., “The Noise and Flow Characteristics of Inverted-Profile 

Coannular Jets,” NASA CR 158995, 1979.

16. Soeder, R., Wnuk, S., and Loew, R. “Aero-Acoustic Propulsion Laboratory Nozzle Acoustic Test 

Rig User Manual,” NASA/TM—2006-212939, 2006.

17. ANSI S1.26-1995 (R2004), “Method for calculation of absorption of sound by the atmosphere”.

18. Amiet, R.K., “Correction of open jet wind tunnel measurements for shear layer refraction,” AIAA77-54, 1977.

TABLE 1.—SETPOINTS FOR HVC EXPERIMENTS

NPRc NPRf NTRc

TTc/Tamb

NTRf

TTf/Tamb

Mfj

1.6000 1.6000 2.9000 1.2900 0.00

1.8000 1.8000 2.9000 1.2900 0.00

1.6000 1.8000 2.6900 1.2900 0.00

1.6000 1.8000 3.0500 1.2000 0.00

1.6000 1.8000 2.9000 1.1000 0.00

1.6000 1.6000 2.9000 1.2900 0.30

1.8000 1.8000 2.9000 1.2900 0.30

1.6000 1.8000 2.6900 1.2900 0.30

1.6000 1.8000 3.0500 1.2000 0.30

1.6000 1.8000 2.9000 1.1000 0.30

Figure 1.—A photograph of the Aero-Acoustic 

Propulsion Laboratory (AAPL) showing the 

Nozzle Acoustic Test Rig (NATR).

NATR

Microphone 

Array
NASA/TM—2012-217744 8

Figure 3.—Cross-section of the GE Inverted Velocity Profile nozzle with 

instantaneous velocity field as predicted by early Large Eddy Simulations

(LES). The LES solutions were provided by GE.

 

Fluid Shield

Hot Flow

Fan Stream

Figure 2.—The models designed by Rolls-Royce Liberty Works for the Supersonics Project in the Fundamental 

Aeronautics Program. The model in (a) is the Highly Variable Cycle (HVC) system, in (b) the N+2 HVC system,

and in (c) the baseline nozzle system for the N+2 HVC system.

Third

Core Stream

Fan Stream

(a) (b)

Sidewall

Mixer Ejector

Sidewall

Fan Nozzle

Ejector Flap

Third-Stream Nozzle

(c)
NASA/TM—2012-217744 9

Figure 5.—The acoustic spectra for the jet conditions in Figure 4 and a simulated forward flight Mach number of 

0.3. The ejector door angles are indicated in the legend. The observation angles are (a) 60o and (b) 160o.

(a) (b)

(a) (b)

Figure 4.—The acoustic spectra acquired at NPRc = 1.60, NPRf = 1.80, NTRc = 2.69, and Mfj = 0.0. The ejector 

door angles are indicated in the legend. The observation angles are (a) 60° and (b) 160°.
NASA/TM—2012-217744 10

(a) (b) (c)

Figure 6.—The results from stereo PIV experiments for cross stream planes located at (a) –0.1D, (b) 

0.7D, and (c) 1.8D relative to the trailing edge of the ejector doors. The effective core-nozzle diameter 

is give by D. The top row contains contour plots of the mean streamwise velocity and cross-stream 

velocity vectors. The bottom row contains contour plots of the turbulent kinetic energy. The conditions 

are the same as those indicated in Figure 4.

m2/s2
NASA/TM—2012-217744 11

(a) (b) (c)

Figure 7.—The results from stereo PIV experiments for cross stream planes located at (a) –0.1D, (b) 0.7D, 

and (c) 1.8D relative to the trailing edge of the ejector doors. The top row contains contour plots of the mean 

streamwise velocity and cross-stream velocity vectors. The bottom row contains contour plots of the 

turbulent kinetic energy. The jet conditions are the same as those indicated in Figure 7 and Mfj = 0.2.
NASA/TM—2012-217744 12

Figure 9.—The acoustic radiation for the N+2 HVC model with a covered ejector flap operated at the conditions 

in Figure 9. The ejector flap angle is 10o and the ejector angle is 9.5°. The observation angles are (a) 90° and 

(b) 150°.

(a) (b)

Figure 8.—The acoustic radiation for the N+2 HVC model operated at core and fan stream conditions 

similar to those in Figure 5 and Mfj = 0.3. The ejector flap angle is 10° and the ejector angle is 9.5°.

The observation angles are (a) 90° and (b) 150°.

(a) (b)
NASA/TM—2012-217744 13

Figure 10.—Streamwise mean velocity contours (top row) and turbulent kinetic energy (bottom row) obtained from 

stereo PIV experiments at cross-stream planes located at (a) 0.1D, (b) 0.4D, (c) 0.8D, (d) 1.6D, and (e) 2.4D relative 

to the trailing edge of the ejector doors. The jet conditions are the same as those in Figure 4 and Mfj = 0.2.

(a) (b) (c) (d) (e)

W m/s
NASA/TM—2012-217744 14

Figure 13.—PIV acquired on the Inverted Velocity Profile nozzle with the fluid shield flow on the bottom half of 

the nozzle. Mean velocity is shown on the left and the rms velocity on the right. Note the low mean velocity 

and high rms turbulent velocity at the separated nozzle lip and the asymmetry introduced by the fluid shield.

Figure 11.—Acoustic results obtained with the GEGR 

model showing the overall sound power levels versus 

ideal thrust for representative low, mid, and high 

temperature cycles with a simulated forward flight Mach 

number equal to 0.3. Also shown are acoustic results 

for reference (REF).

Figure 12.—Instantaneous velocities obtained 

from the PIV experiments showing separation 

on the nozzle lip between inner and outer 

streams.
NASA/TM—2012-217744 15

(a) (b)

Figure 15.—The acoustic spectra for the IVP and reference nozzle acquired at observation angles equal 

to (a) 90° and (b) 150°.

Figure 14.—The overall sound pressure level 

directivity for the IVP and reference nozzle.
REPORT DOCUMENTATION PAGE Form Approved 

OMB No. 0704-0188 

The public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources, gathering and maintaining the 

data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this collection of information, including suggestions for reducing this 

burden, to Department of Defense, Washington Headquarters Services, Directorate for Information Operations and Reports (0704-0188), 1215 Jefferson Davis Highway, Suite 1204, Arlington, VA 22202-4302. 

Respondents should be aware that notwithstanding any other provision of law, no person shall be subject to any penalty for failing to comply with a collection of information if it does not display a currently valid OMB 

control number. 

PLEASE DO NOT RETURN YOUR FORM TO THE ABOVE ADDRESS. 

1. REPORT DATE (DD-MM-YYYY)

01-11-2012 

2. REPORT TYPE

Technical Memorandum 

3. DATES COVERED (From - To)

4. TITLE AND SUBTITLE

Jet Noise Reduction Potential From Emerging Variable Cycle Technologies 

5a. CONTRACT NUMBER

5b. GRANT NUMBER

5c. PROGRAM ELEMENT NUMBER

6. AUTHOR(S)

Henderson, Brenda, S.; Bridges, James, E.; Wernet, Mark, P. 

5d. PROJECT NUMBER

5e. TASK NUMBER

5f. WORK UNIT NUMBER

WBS 984754.02.07.03.17.08.01 

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration 

John H. Glenn Research Center at Lewis Field 

Cleveland, Ohio 44135-3191 

8. PERFORMING ORGANIZATION

 REPORT NUMBER

E-18494 

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration 

Washington, DC 20546-0001 

10. SPONSORING/MONITOR'S

 ACRONYM(S)

NASA 

11. SPONSORING/MONITORING

 REPORT NUMBER

NASA/TM-2012-217744 

12. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified-Unlimited 

Subject Category: 71 

Available electronically at http://www.sti.nasa.gov 

This publication is available from the NASA Center for AeroSpace Information, 443-757-5802

13. SUPPLEMENTARY NOTES

14. ABSTRACT

Acoustic and flow-field experiments were conducted on exhaust concepts for the next generation supersonic, commercial aircraft. The 

concepts were developed by Lockheed Martin (LM), Rolls-Royce Liberty Works (RRLW), and General Electric Global Research (GEGR) 

as part of an N+2 (next generation forward) aircraft system study initiated by the Supersonics Project in NASA’s Fundamental Aeronautics 

Program. The experiments were conducted in the Aero-Acoustic Propulsion Laboratory at the NASA Glenn Research Center. The exhaust 

concepts utilized ejectors, inverted velocity profiles, and fluidic shields. One of the ejector concepts was found to produce stagnant flow 

within the ejector and the other ejector concept produced discrete-frequency tones that degraded the acoustic performance of the model. The 

concept incorporating an inverted velocity profile and fluid shield produced overall-sound-pressure-level reductions of 6 dB relative to a 

single stream nozzle at the peak jet noise angle for some nozzle pressure ratios. Flow separations in the nozzle degraded the acoustic 

performance of the inverted velocity profile model at low nozzle pressure ratios.

15. SUBJECT TERMS

Jet noise 

16. SECURITY CLASSIFICATION OF: 17. LIMITATION OF

 ABSTRACT

UU 

18. NUMBER

 OF 

 PAGES

22 

19a. NAME OF RESPONSIBLE PERSON

STI Help Desk (email:help@sti.nasa.gov) 

a. REPORT

U 

b. ABSTRACT

U 

c. THIS 

PAGE

U 

19b. TELEPHONE NUMBER (include area code)

443-757-5802 

Standard Form 298 (Rev. 8-98)

Prescribed by ANSI Std. Z39-18