Dale E. Van Zante and Edmane Envia

Glenn Research Center, Cleveland, Ohio

Prediction of the Aero-Acoustic Performance 

of Open Rotors

NASA/TM—2014-218132

December 2014

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Dale E. Van Zante and Edmane Envia

Glenn Research Center, Cleveland, Ohio

Prediction of the Aero-Acoustic Performance 

of Open Rotors

NASA/TM—2014-218132

December 2014

GT2014–26413

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NASA/TM—2014-218132 1 

PREDICTION OF THE AERO-ACOUSTIC PERFORMANCE OF OPEN ROTORS 

Dale E. Van Zante and Edmane Envia 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

ABSTRACT 

The rising cost of jet fuel has renewed interest in contrarotating open rotor propulsion systems. Contemporary design 

methods offer the potential to maintain the inherently high 

aerodynamic efficiency of open rotors while greatly reducing 

their noise output, something that was not feasible in the 1980’s 

designs. The primary source mechanisms of open rotor noise 

generation are thought to be the front rotor wake and tip vortex 

interacting with the aft rotor. In this paper, advanced 

measurement techniques and high-fidelity prediction tools are 

used to gain insight into the relative importance of the 

contributions to the open rotor noise signature of the front rotor 

wake and rotor tip vortex. The measurements include threedimensional particle image velocimetry of the intra-rotor 

flowfield and the acoustic field of a model-scale open rotor. The 

predictions provide the unsteady flowfield and the associated 

acoustic field. The results suggest that while the front rotor tip 

vortex can have a significant influence on the blade passing 

tone noise produced by the aft rotor, the front rotor wake plays 

the decisive role in the generation of the interaction noise 

produced as a result of the unsteady aerodynamic interaction of 

the two rotors. At operating conditions typical of takeoff and 

landing operations, the interaction noise level is easily on par 

with that generated by the individual rotors, and in some cases 

is even higher. This suggests that a comprehensive approach to 

reducing open rotor noise should include techniques for 

mitigating the wake of the front rotor as well as eliminating the 

interaction of the front rotor tip vortex with the aft rotor blade 

tip. 

INTRODUCTION 

Open Rotors demonstrated significant fuel burn advantage 

compared to turbofans in 1980’s era ground and flight tests (1,

2). Blade designers of the time were not able to optimize for 

both high aerodynamic efficiency and low noise due to 

limitations of the design and analysis tools of that time. This led 

to compromises in aerodynamic efficiency in order to meet the 

noise regulation requirements for a product. 

Current open rotor efforts make extensive use of modern 

three-dimensional computational fluid dynamics (CFD) and 

computational aeroacoustics (CAA) tools to produce designs 

that are both high-efficiency and low-noise (3-5). Additionally, 

computational methods have been used to further investigate the 

mechanisms of open rotor noise generation in order to develop 

rules for low-noise blade designs. Much of this type of work has 

employed notional blade designs and blade counts, since 

realistic blade geometries tend to be highly proprietary and not 

available to the public. Recent examples of this type of work 

include Danner, et. al (6), where unsteady CFD simulations 

were analyzed to identify specific aerodynamic mechanisms that 

lead to open rotor noise generation. Peters and Spakovszky (7) 

used aero and acoustic simulation tools to associate various 

blade regions with specific noise generation mechanisms, 

though their choices of blade regions are somewhat arbitrary. In 

addition to establishing the important role of 3D blade 

aerodynamics, results from these analyses confirmed already 

established conclusions that rotor-rotor spacing and aft rotor tip 

clipping are effective noise reduction tools. As noted by 

Brandvik, et al (8), angle of attack effects influence the 

optimum tip clipping and rotor spacing. 

In contrast, the work presented here uses realistic highperformance blade geometries with blade counts that are 

representative of modern open rotor designs. Furthermore, the 

acoustic source identification methodology used in the present 

work does not rely on a priori assumptions for associating 

particular region(s) of the blade to specific noise source(s). 

The experimental data used here to establish the validity of 

the theoretical analysis were acquired as part of a 

comprehensive research effort funded by the NASA 

Environmentally Responsible Aviation (ERA) Project. Under 

the sponsorship of the ERA project, extensive datasets of 

flowfield and acoustic measurements were acquired for a 

benchmark blade design called F31/A31 to enable validation 
NASA/TM—2014-218132 2 

and improvement of design/analysis codes for contra-rotating 

open rotors (9). An important element of that research program 

was the acquisition of three-component velocity measurements 

in the intra-rotor region (i.e., the region between the two rotors) 

using stereo particle image velocimetry (PIV). Previously, this 

PIV dataset was used to analyze the front rotor tip vortex 

trajectory (10). In the present work, the PIV dataset is used to 

validate flowfield simulation results generated by the 

NUMECA’s FINETM/Turbo code. The validated CFD 

computations are then used as input to an acoustic model to 

predict tone noise produced by the open rotor, which, in turn, is 

compared with the measured noise results. The goal of this 

study is to gain insight into the noise generating regions of the 

blades and to suggest potential noise mitigation approaches for 

reducing open rotor noise beyond what has been achieved 

already. 

NOMENCLATURE 

Mach Test section free stream Mach number 

β Blade setting angle at 75% radius (degrees) 

Subscripts: 

1, 2 Front, Aft blade 

C Standard Day Corrected 

EXPERIMENTAL CONFIGURATIONS 

The wind tunnel configurations and the blade set used for 

the acquisition of aerodynamic, acoustic, and detailed flowfield 

measurements have been described in detail in Refs. 10-12. An 

overview of the configurations is provided here for the sake of 

completeness. 

Open Rotor Testbed 

The so-called Historical Baseline blade set, F31/A31, is 

used for all measurements presented in this paper. F31/A31 is 

representative of early 1990s aero design technology and is 

optimized for performance without any compromises for 

lowering its acoustic signature. The front and aft rotor blade 

counts for F31/A31 are 12 and 10, respectively. The design tip 

speed of the rotors is 230 m/s (750 ft/s). A side view of the front 

and the aft blades is shown in Figure 1. Geometric parameters 

are given in Table 1. 

Table 1: F31/A31 Geometric Parameters 

F31 Diameter 0.652 m (25.662 inches) 

A31 Diameter 0.630 m (24.794 inches) 

Diameter Ratio 

(Aft Rotor/Front Rotor) 

0.966 

Front Rotor Blade Count 12 

Aft Rotor Blade Count 10 

Blade Spacing to 

Front Rotor Diameter Ratio 

0.31 

The F31/A31 aerodynamic and acoustic measurements 

were acquired in the NASA Glenn Research Center 9-foot by 

15-foot Low Speed Wind Tunnel (9x15 LSWT) using the 

refurbished Open Rotor Propulsion Rig (ORPR). Figure 2 

shows the ORPR installed in the 9x15 test section with the 

traversing microphone system that was the primary acoustic 

acquisition tool for this test. Test section freestream conditions 

were set using a rake mounted on the test section ceiling at an 

axial location near the rotor plane. Data presented here are 

standard day corrected to the test section static temperature as is 

the convention for propellers. 

Figure 1: The Historical Baseline blade set, F31/A31. 

The primary rotor performance instrumentation was 

rotating force balances in each rotor hub. The maximum force 

balance capabilities were ~1,910 N (430 lbf) of thrust and ~745 

N-m (550 ft-lbf) of torque. The thrust and torque values are 

accurate to within 1.0% of force level as determined from check 

loads of the balances installed in the rotor hubs on the rig. 

Additionally, the rotor balance thrust values were corrected for 

pressure-area forces on the rotor disks beneath the flow path 

and tare drag on the hub contour as determined from 

measurements with no rotor blades. Details are given in Refs. 5 

and 11. Signals were transmitted from the rotating frame to the 

stationary reference frame with a digital telemetry system. 

Acoustic measurements were acquired at a sideline distance 

of 1.524 m (5 ft) at 18 stops along the traverse track which 

correspond to 18 sideline geometric angles spanning a range 

from roughly 18 degrees to 140 degrees with 0 degrees 

denoting the upstream direction and 180 degrees the 

downstream direction. The microphone signal was digitized at 

200 kHz for 15 seconds per stop. Spectra were then generated 

using a 214 point FFT resulting in a frequency bin width of 12.2 

Hz. Details of the acoustic processing, instrument corrections, 

atmospheric corrections, etc. are discussed in Ref. 12. The 

accuracy of the acoustic measurement system is ±1 dB. 
NASA/TM—2014-218132 3 

Performance and acoustic data were acquired at two blade 

setting angles corresponding to the nominal takeoff (NTO) and 

approach conditions, at several rotor speeds, multiple angles of 

attack, and at two different test section Mach numbers. For the 

purposes of this paper, the focus is on a subset of the test data 

corresponding to the NTO blade setting angles, six tip speeds, 

zero angle of attack, and one Tunnel Mach number. The blade 

setting angles considered are β1 = 40.1 degrees and β2 = 40.8 

degrees. The range of speeds is listed in Table 2 and includes 

the nominal design tip speed of 6436 RPM at NTO down to a 

low part speed of 4620 RPM. Note that the test matrix 

considered here includes only the equal RPM cases. 

Furthermore, the RPMs indicated are corrected speeds. 

Table 2. Rotor RPMs Considered in This Study 

Case Rotor 1 RPM Rotor 2 RPM 

1 6436 6436 

2 6303 6303 

3 6068 6068 

4 5551 5551 

5 5268 5268 

6 4620 4620 

Figure 2: The F31/A31 Open Rotor installed in 9x15 

LSWT. Traversing microphone and its track are seen 

on the left side. 

Stereo PIV Setup 

Three-component velocities were measured using the 

stereo PIV technique for the intra-rotor region. To permit an 

unobstructed view, the lasers were mounted above the test 

section and the cameras were in the test section wall as shown 

in Figure 3. The cameras and lasers were set on three 

independent translation tracks to measure a series of planes 

from near the hub to the outboard region of the rotor tips. The 

light sheet was clipped in the axial dimension at the lower spans 

by a motorized aperture. The radial spacing of the planes was 

variable with increased resolution near the blade tip. Figures 4 

and 5 show the data volume and measurement locations. The 

flow was seeded far upstream of the test section using Vicount 

smoke generators which generated condensed mineral oil 

droplets in the range of 0.2-0.3 μm in diameter at high volumes. 

There was no noticeable buildup of mineral oil on the blade 

surfaces during the elapsed time to acquire the PIV data. 

Extensive PIV system calibrations were performed and 

multi-step image processing applied to obtain the highest 

quality and best spatial resolution velocity vector maps 

possible. See Ref. 10 for details of the PIV processing. To keep 

the total data acquisition time reasonable, the measurements 

were synchronized to the forward rotor position only. That is 

when any of the forward rotor blades was in the correct 

position, and the laser/camera system was ready to fire, an 

image pair would be acquired. Blade to blade differences in 

front rotor geometry or blade setting angle thus appear as 

‘turbulence’ in the ensemble averages. The aft rotor position is 

arbitrary, thus it is assumed that the upstream influence of the 

aft rotor potential field is minimal. The aft rotor potential field 

influence would also appear as ‘turbulence’ in the ensemble 

average velocity field. This point will be discussed further in 

the Intra-Rotor Velocity Field Comparisons section below. 

laser

Camera 1

Camera 2

Laser

sheet

Figure 3: Stereo PIV configuration in the 9x15 LSWT 

with the Open Rotor Propulsion Rig. 

Sequences of 400 velocity vector maps were acquired at 

each measurement station and were ensemble-averaged to 

provide first and second order statistics over the entire 

measurement plane. The wind tunnel has very low free stream 

turbulence and thus experience has shown that an average of 

400 vector maps is adequate to converge the velocity statistics 
NASA/TM—2014-218132 4 

given the other potential sources of uncertainty. All of the 

processed PIV data were placed in the model coordinate system 

to facilitate comparison with CFD predictions. The final 

processed velocity vector maps had an in-plane spatial 

resolution of 1.43 mm in both directions in the measurement 

plane. The velocity measurements have an accuracy of better 

than 1% of full scale. 

Figure 4: Perspective view of the PIV intra-rotor 

measurement region. 

Figure 5: PIV measurement planes (a total of 30 

horizontal planes). Note that the spanwise and 

streamwise extent of the measurement planes get 

smaller as one moves down the span towards the 

hub. 

The PIV data were acquired for multiple blade setting 

angles and speeds, but the focus here is on the one condition 

corresponding to the NTO at 6303 RPM, which is highlighted 

in Table 2. 

COMPUTATIONAL SETUP 

The computational tool used for the aerodynamic 

calculations in this work is the commercial CFD software 

package FINETM/Turbo developed by NUMECA International. 

FINETM/Turbo is a turbomachinery CFD simulation software 

package with integrated meshing and post-simulation analysis 

tools. It is a structured, multi-block, unsteady Navier-Stokes 

solver, which offers several solution algorithm choices along 

with several acceleration strategies. To reduce the 

computational time requirements for the time-dependent 

simulations, FINETM/ Turbo was run in the nonlinear harmonic 

(NLH) mode which solves for a finite number of the blade 

passing frequency harmonic components of the time-dependent 

solution, but ignores all the other unsteady components (13). 

For this study, only three loading harmonics were retained to 

keep the computational cost reasonable. This implies that tones 

up to the 66th shaft order could be modeled. 

The FINETM/Turbo computational domain used in 

generating the results presented in this paper includes one 

passage each of the two blade rows and their associated 

ancillary domains like the spinner, hub, farfield, etc. The total 

mesh size is slightly over 27.1 million grid points with the 

farfield boundary set seven tip radii away. The equivalent fullwheel grid for a full unsteady simulation would require almost 

300 million grid points. The Spalart-Almaras turbulence model 

was used in this simulation to best represent the physics of the 

problem. Further details of the simulation setup are given in 

Ref. 14. The computational domain is shown in Figure 6. 

Figure 6: CFD computational domain, which contains 

91 computational blocks and 27.1 million grid points. 
NASA/TM—2014-218132 5 

The acoustic computations were carried out using a NASA 

code called LINPROP, which is based on a high-blade-count 

asymptotic approximation of the Ffowcs Williams-Hawkings 

equation (see Ref. 14 for further detail). The LINPROP code 

was used to predict the tone noise on a 5-foot (1.524 m) sideline 

parallel to the rotational axis of the open rotor. This duplicates 

the experimental setup in the wind tunnel where the acoustic 

data were acquired using the traversing microphone. The 

nacelle geometry, rotor geometry and aerodynamic pressure 

distributions from the CFD calculation are direct inputs to the 

LINPROP calculation. 

PERFORMANCE AND FLOW FIELD COMPARISONS 

To evaluate the accuracy of the CFD simulations, overall 

performance metrics, such as rotor thrust and torque ratio (i.e., 

ratio of aft rotor torque to the front rotor torque), are first 

compared to the numerical results and then more detailed 

comparisons of the intra-rotor velocity fields follow. 

Overall Performance Comparisons 

FINETM/Turbo aerodynamic simulations were generated for 

six rotor speeds that match, on corrected RPM basis, a subset of 

the conditions from the wind tunnel test. Figure 7 shows the 

data-prediction comparisons for the total thrust and torque ratio 

at the corrected speed of 6303 RPMC. This is the corrected 

speed for which the PIV data were acquired. The overall 

predicted thrust values match to better than 1.5% with the 

measured data. The torque ratio values match to better than 

11%. It should be noted that the FINETM/Turbo calculations 

were true predictions in the sense that no attempt was made to 

modify the grid, boundary conditions, or operating condition to 

match a measured parameter. 

Figure 7.Comparison of measured and predicted 

total rotor thrust and torque ratio. 

Furthermore, the F31/A31 rotor system geometry, which 

was supplied by GE, is for the top of climb condition. For the 

purposes of the work presented here, the blades were re-pitched 

to match the nominal takeoff blade setting angles, but no 

attempt was made to compute the hot shapes corresponding to 

the NTO condition. It was expected that the blade shape 

differences would be relatively small at least for the condition 

considered in this paper. Next, the details of the velocity field 

are compared for the 6303 RPMC condition. 

Table 3: Performance Comparison 

Data vs. Simulation at PIV Condition 

RPMC Thrust, N (lbf) Torque Ratio 

6303 Measured Predicted Measured Predicted 

2539 (571) 2581 (580) 0.962 1.069 

Intra-Rotor Velocity Field Comparisons 

The distribution of the axial velocity field is compared at 

two axial locations in the intra-rotor region as shown in Figure 

8. For reference, the aft rotor pitch change axis is at the origin 

(i.e., 0.0 cm). The forward rotor pitch change axis is at -19.91 

cm (-7.84 in). Velocity comparisons are shown at -13.30 cm 

(-5.25 in) and at -5.70 cm (-2.25 in). The first location is close 

to the forward rotor trailing edge in order to compare the front 

rotor wake shape and tip vortex character. The second location 

is close to the aft rotor leading edge and shows how the wake 

and tip vortex have evolved before striking the aft rotor. 

Figure 8: Locations for detailed intra-rotor velocity 

comparisons. 

The predicted axial velocity fields for the full 

circumference at these two axial planes are shown in Figure 9. 

Unlike the ensemble-averaged PIV results to be discussed later, 

the simulated velocity fields are for an instant in time. Figure 

9A shows essentially a periodic pattern of the forward rotor 

wakes and tip vortices 12 times around the circumference. The 

distribution shows almost no passage-to-passage variations, 
NASA/TM—2014-218132 6 

which implies that the aft rotor potential field influence is quite 

weak at this location. This conclusion bears out the assumption 

made in acquiring the PIV data. However, the downstream 

station distribution, shown Figure 9B, exhibits significant 

circumferential variation due to the 12 wake/vortex pattern of 

the front rotor interacting with the potential field of the 10-

bladed aft rotor. The forward rotor wakes are visible as 12 low 

velocity regions at the inner diameter. However, the image is 

dominated by 10 higher velocity regions, which are due to the 

potential field influence of the aft rotor. 

In light of these results, the interpretation of the PIV data is 

crucial for meaningful data-theory comparisons. In the planes 

near the front rotor, it is expected that the ensemble-averaged 

PIV maps are effectively equivalent to the instantaneous 

velocity distributions like that shown in Figure 9A since there 

are virtually no passage-to-passage variations in the flowfield at 

these locations. On the other hand, for the planes close to the aft 

rotor, the PIV maps should be noticeably different from the 

instantaneous picture like that shown in Figure 9B. That is 

because the relative position of the front and aft rotor blades is 

not known at the instant that any of the PIV images was 

acquired. As such the ensemble average is not phased-locked to 

any known fixed relative positions of the two rotors. Therefore, 

since it is not possible to ensemble-average the CFD results in 

precisely the same manner that the PIV averages were 

constructed, we carry out a straightforward front rotor phaselocked averaging as an approximation for the sake of 

comparisons with the PIV data near the aft rotor. This is an 

average using six passages. Note that given the 12/10 blade 

counts and the equal front and aft rotor RPMs, the simulated 

flowfield is periodic in a half-wheel so only half the field needs 

to be considered in ensemble averaging. For the sake of 

illustrating the instantaneous variations, some representative 

ones together with the ensemble average are shown. 

Figure 10 shows the velocity comparison at the front axial 

location. The PIV results show a clearly defined tip vortex in 

the upper center of the velocity field. Two forward rotor blade 

wakes are also visible in the velocity field. The CFD results at 

the same axial plane show remarkably similar features. Figure 

11A shows axial velocity along a radial line that passes through 

the velocity minimum of the tip vortex in the PIV results (the 

white dashed line). The identical location was extracted from 

the CFD solution and plotted for comparison. Note the excellent 

agreement in the core flow velocity. The PIV results show a 

larger velocity difference through the vortex. Figure 11B shows 

a comparison of wake profiles on a constant radius cut (the 

white curved dashed line). Again, the core flow velocities and 

wake widths show good agreement. The CFD shows a deeper 

wake. Overall agreement of absolute values, location and 

characteristics of flow non-uniformities is quite good at this 

axial location. 

The data/CFD comparison at the downstream location, 

shown in Figure 12, is more complex. The PIV data in Figure 

12A show a diminished, though still distinct, tip vortex and 

forward rotor blade wakes. Note that the radial extent of the 

PIV data at that station is limited (recall Figure 5) in order to 

mitigate the risk of the laser sheet damaging the aft rotor blade. 

The CFD flow features (Figure 12B) are less distinct. A more 

revealing comparison is shown in Figure 13, where the CFD 

results are also averaged though not exactly in the same manner 

as the PIV data as discussed earlier. Detailed comparisons along 

a radial line and a constant radius are shown in Figure 13. Both 

A. Axial location -13.30 cm (-5.25 in) B. Axial location -5.70 cm (-2.25 in) 

Figure 9: Instantaneous axial velocity (W) contours in two axial planes extracted from the CFD solutions. 

The left pane shows the results for the axial plane close to the front rotor trailing edge and the right pane 

shows the results for the plane close to the leading edge of the aft rotor. 
NASA/TM—2014-218132 7 

the tip vortex (Figure 13A) and blade wakes (Figure 13B) are 

more mixed out in the CFD solution compared to the ensemble 

average PIV data. The three instantaneous velocities from the 

CFD solution are also shown as dashed lines. The solid magenta 

lines show the phase-locked average of the CFD results to serve 

as an approximation to the way the PIV data were generated. 

While, for the reasons discussed earlier, the comparison would 

never be exact, it is nonetheless clear that the predictions 

exhibit more dissipation than is seen in the measured data. 

However, given the limitations of this comparison, and the fact 

that the data-theory difference is actually small, i.e., about 5% 

in the maximum wake velocity deficit (see Figure 13B), the 

agreement is still quite reasonable. In any case, small 

differences do not have a large impact on the acoustic results, 

since a 10% change in the aerodynamic input level translates 

into a 1 dB change in the acoustic level. 

A. PIV measurements B. CFD simulations 

Figure 10: Measured and predicted axial velocity (W) contours in a plane 13.36 cm (5.26 inches) upstream 

of the aft rotor pitch change axis. Dashed lines show locations of velocity profiles depicted in Figure 11. 

A. Velocity along a radial line through the tip vortex B. Wake profile on a constant radius 

Figure 11: Comparison of measured and predicted axial velocity (W) in a plane 13.36 cm (5.26 inches) 

upstream of the aft rotor pitch change axis. 
NASA/TM—2014-218132 8 

A. PIV measurements B. CFD simulation results 

Figure 12: Measured and predicted axial velocity (W) contours in a plane 5.70 cm (2.25 inches) upstream 

of the aft rotor pitch change axis. Dashed lines show locations of velocity profiles depicted in Figure 13. 

A. Velocity along a radial line through the tip vortex B. Wake profiles on a constant radius 

Figure 13: Comparison of measured and predicted axial velocity (W) in a plane 5.70 cm (2.25 inches) 

upstream of the aft rotor pitch change axis. 
NASA/TM—2014-218132 9 

ACOUSTIC RESULTS

The blade harmonic loading distributions extracted from 

the CFD simulations were used as input to the LINPROP code 

in order to predict the tone spectra of F31/A31 at various 

conditions. Both individual rotor tones (denoted by nBPF1 and 

nBPF2) as well as interaction tones (denoted by mBPF1 + 

kBPF2) were computed on the 1.526 meters sideline at a large 

number of angles encompassing all of the 18 acoustic 

measurements locations used in the wind tunnel test. Note that 

m, n and k in the above expressions are arbitrary integers. 

Since open rotor acoustic spectra are rich in tone content, it 

is useful for the subsequent discussion to show an example of 

the measured spectra and identify some of the important aspects 

of the spectrum. Figure 14 shows the measured sound pressure 

level (SPL) spectrum for F31/A31 at the 6303 RPMC condition 

at the broadside angle of 90 degrees. The abscissa is given in 

terms of the rotor shaft orders, since it makes the tone 

identification easier. Recall that F31/A31 has 12 front blades 

and 10 aft blades and the rotor rotational speeds are equal for 

the cases considered here. It is typical to have open rotor tones 

rising well above the broadband level up to the 200th shaft 

order, but for the sake of clarity only tones up to the 48th shaft 

order are shown in Figure 14. Tones below the 8th shaft order 

are contaminated by the rig noise or tunnel background noise 

and are not shown. The labels identify some of the expected 

individual rotor tones and interaction tones based on the theory. 

However, clearly there are more tones in the measured spectrum 

than indicated by the theory. The preponderance of tones in the 

measured spectrum is due to the fact that the blades in each 

rotor are never identical nor can they ever be set in the hub with 

precisely the same blade setting angles. These imperfections 

result in the scattering of some of the acoustic energy into shaft 

orders other than those that would be predicted if the blades in 

each rotor were identical in every respect. Even small variations 

in the blades in each rotor destroy the perfect phase relationship 

assumed by the theory and result in the generation of extraneous 

tones. This inherent discrepancy should be kept in mind when 

comparing the theoretical predictions to the measured data. It 

should also be noted that the number of interaction tones in the 

spectrum far exceeds that of the individual rotor tones. In fact, 

for takeoff and approach conditions, the interaction tones levels 

can be noticeably higher than those of individual rotor tones, a 

point that is germane to the main thrust of this paper. The 

importance of this observation will be discussed shortly. 

Figure 15 shows the comparison of measured and predicted 

SPL spectra for the blade passing tones of the front and aft 

rotors (i.e., BPF1 & BPF2) as well as the first interaction tone 

(i.e., BPF1 + BPF2) over a range of sideline angles. It should be 

noted the predicted tone levels are confined to a bin width of 

zero Hertz while the measured tone levels are spread across 

multiple frequency bins. Therefore, in order to compare the 

measured levels to the predicted ones, it was necessary to sum 

the tone energy in the bins over which the measured tone is 

spread. Furthermore, where the tone level is close to the 

broadband level (i.e., less than 6 dB), tone extraction is 

problematic. That is because it is not clear how much of the 

energy is due to the tone and how much of it is associated with 

the underlying broadband level. As such, where the measured 

tone level is less than 6 dB above the background level, the 

measured tone level is not plotted. The results in Figure 15 

show that the absolute tone levels are not well predicted by the 

theory primarily as a result of the idealization assumed in both 

the CFD simulations and the LINPROP acoustic calculations as 

was discussed earlier. Nevertheless, the trends with the sideline 

angle are fairly well predicted by the theory. In fact, predicting 

acoustic trends is often a more useful tool in guiding the design 

than the absolute levels. Therefore, the authors felt justified in 

using the LINPROP code to analyze and ultimately identify the 

dominant noise generating regions of the blade that control the 

radiated acoustic field. Before reviewing the results of the 

analysis, it is important to elucidate certain aspects of the results 

in Figure 15. 

First, note the clear difference between the directivities of 

the individual rotor tones and the interaction tones. Whereas the 

individual rotor tones tend to peak around angles centered on 

the broadside (i.e., 90 degrees), the interaction tones tend to 

radiate substantially more towards the upstream and/or 

downstream directions as compared with the broadside 

direction. This is seen in the measured data as well as the 

predictions. It is important to point out that the source of the 

individual rotor tones is the steady loading on the rotors blades, 

whereas the source of the interaction tones is the unsteady 

loading induced on the blades as a result of the aerodynamic 

Figure 14: Comparison of absolute tone levels for 

the first 48 shaft orders at 90 degrees, (i.e., 

broadside) to the aft rotor for the 6303 RPMC speed. 

A few individual rotor tones and some interaction 

tones are highlighted. The spectrum below the 8th 

shaft order is contaminated by rig noise and tunnel 

background noise. 
NASA/TM—2014-218132 10 

coupling between the two rotors. Interestingly, though the 

steady loading is at least two orders of magnitude larger than 

the perturbation loading, the substantially higher radiation 

efficiency of the interaction tones (see Ref. 14) more than 

compensates for their small source strength thus generating 

levels that for some interaction tones exceed the individual rotor 

tones as seen in the example results shown in Figure 14. 

Figure 15: Data-Theory comparisons for the first 

three principal tones: top BPF2, middle BPF1, and 

bottom BPF1+BPF2. Solid lines denote predictions 

from LINPROP and symbols denote wind tunnel 

data. Data have a ±1dB uncertainty band. 

Another noteworthy aspect of the interaction tones is that 

while both rotors can generate any given interaction tone, the 

theoretical analysis indicates that the aft rotor contribution tends 

to be far more substantial than that of the front rotor, typically 

by a factor of 10. The reason for this behavior is related to the 

fact that the viscous coupling of the two rotors is far stronger 

than their potential field coupling as attested to by the 

discussion of the results in Figure 9. The analysis also indicates 

that the aft rotor harmonic tones (i.e., nBPF2) are stronger than 

those of the front rotor (i.e., nBPF1). Taken together these 

observations substantiate the widely held belief that the aft rotor 

is the dominant source of open rotor noise and as such has been 

the focus of noise reduction research. 

A substantial portion of open rotor noise reduction research 

has focused on the tip vortex as the main culprit of aft rotor 

noise generation. In that vein, the most popular noise reduction 

approach has been to clip the aft rotor blade tip and/or to 

modify the front rotor blade tip so as to mitigate the strength of 

the tip vortex interaction with the aft blade. While this approach 

has proven successful in reducing open rotor noise, a more 

potent approach must also consider the importance of the front 

rotor wake impingement on the aft rotor. That is because the 

wake not only substantially influences the aft rotor steady 

loading, it also controls its unsteady loading. As such it has a 

significant influence on the aft rotor tones as well as the 

interaction tones. As was mentioned in the introduction, 

attempts have been made in past to quantify the relative 

importance of the contributions from the tip vortex and blade 

wake to the overall noise of the open rotor (e.g., Ref. 7). 

However, these analyses have been somewhat ad hoc in their 

choices for assessing the relative importance of the tip vortex 

versus the blade wake. 
NASA/TM—2014-218132 11 

In contrast, in the present study, the question of the relative 

importance of the tip vortex and blade wake has been addressed 

by a systematic analysis of the contributions to the noise field 

from the aft rotor blade span without any á priori choices about 

which portion of the blade may be important for a given source. 

The results from this analysis are summarized in Figures 16 

through 19 for select (but representative) aft rotor tones and 

some important interaction tones. The selected tones are BPF2, 

2BPF2, BPF1+BPF2, BPF1+2BPF2. Each figure includes a pair 

of plots with the top portion showing the predicted SPL sideline 

directivity of the tone and the bottom portion a contour plot of 

the contribution to the sideline SPL directivity from the various 

regions on the blade. Specifically, the contour shows the change 

in tone level as a function of aft rotor blade percent span that is 

included in the noise calculations. The contour level is 

referenced to the full-span tone level, so that the relative SPL 

(i.e., ΔSPL) is zero when the entire span is taken into account. 

Figure 16: Change in BPF2 tone SPL as a function of 

aft rotor blade percent span included in noise 

calculations relative to the full-span tone level. 

It should be noted that the partial blade loading levels have 

been corrected for the loss of thrust that would be experienced 

with a partial blade by raising the loading level on the partial 

blade by multiplying the loads by an amount equal to the thrust 

ratio Tfull span / Tpart span, where Tfull span is the aft rotor thrust when 

the entire span is taken into account, and Tpart span is when only 

part span is considered. This has the effect of increasing the 

affected tone dB level by 10 log10 (Tfull span / Tpart span). The 

correction is negligible near the tip, but is around 16 dB when 

very small portions of the span are considered. The correction 

does not change the behavior observed in the contour plots, but 

has the effect of bunching up the contours. Finally, since only 

the steady loads contribute to blade thrust, no correction has 

been applied to the unsteady loading. As a result, only the levels 

of rotor tones have been adjusted, but not the interaction tone 

levels. 
NASA/TM—2014-218132 12 

The contour levels shown in Figure 16 through 19 should 

then be interpreted as follows. Take Figure 16 for example. If 

one takes into account the contributions from the first 75% of 

the aft blade span to the BPF2 tone level, one finds that the 

resulting sideline noise level will be approximately 10 dB less 

than the full span level across the entire sideline angle range. 

Therefore, it quickly becomes clear that the last 25% or so of 

the aft blade span contributes about 90% of the radiated level of 

the BPF2 tone since dB is a logarithmic scale. An even smaller 

portion of the outboard section of the blade is needed (about 

10%) to generate most of the 2BPF2 tone level on the sideline 

as see in Figure 17. By contrast, the inboard region of the blade 

Figure 18: Change in BPF1+BPF2 tone SPL as a 

function of aft rotor blade percent span included in 

noise calculations relative to the full-span tone level. 

contributes only a minute fraction (about 1%) to the levels of 

the BPF2 and 2BPF2 tones below about 60% of the span. The 

same story holds for the higher harmonics of the BPF2 tone. 

This result would appear to substantiate the view that clipping 

the aft rotor tip (say, to about 80% span) should have a 

measurable affect on reducing the total open rotor tone noise (in 

this case about 10 dB for BPF2 tone level and 15 dB for 2BPF2

tone). However, this would be misleading. 

As it can be clearly seen in Figures 18 and 19, the behavior 

of the interaction tones BPF1+BPF2 and BPF1+2BPF2 is 

radically different. Instead of the outboard portion of the blade 

being the main source of noise, the entire span is involved in 

Figure 17: Change in 2BPF2 tone SPL as a function of 

aft rotor blade percent span included in noise 

calculations relative to the full-span tone level. 
NASA/TM—2014-218132 13 

Figure 19: Change in BPF1+2BPF2 tone SPL as a 

function of aft rotor blade percent span included in 

noise calculations relative to the full-span tone level. 

generating the tone level as evidenced by the complex ΔSPL

contour patterns for these tones. In fact, given positive contour 

levels associated with some partial blade spans, it becomes 

clear that the radiated acoustic field of an interaction tone is a 

complex amalgam of the contributions from the entire span with 

no one particular region holding sway over the tone level. In 

fact, if the blade is clipped injudiciously, the interaction tone 

level could actually increase owing to reduced interference 

caused by the missing blade portion. This is most clear for the 

BPF1+2BPF2 tone between 80 and 100 degrees on the sideline. 

For example, if the blade were clipped by 20%, the tone level 

would go up by about 5 dB. Given the measured tone levels 

shown in Figure 14, such clipping would result in a substantial 

net increase in the noise because, while the levels of the BPF2

and 2BPF2 tones would decrease, the increase in the level of 

BPF1+2BPF2 tone would more than offset these benefits. 

To further illustrate the importance of the interaction tones 

at operating conditions like takeoff and approach, in Figure 20 

the predicted sideline directivity of the overall sound pressure 

level (OASPL) for all the aft rotor tones, i.e., ∑ nBPF2, and the 

OASPL for all the interaction tones, i.e., ∑ mBPF1 + kBPF2, are 

plotted for various tip speed conditions investigated in this 

study. The OASPL for the aft rotor tones (shown in black) 

includes five harmonics, but the BPF2 level dominates the total. 

For the interaction tones (shown in orange), the OASPL 

includes nine tones, which are all the interaction tones that 

could be calculated given that the aerodynamic simulations had 

only three loading perturbations included (see the discussion in 

the COMPUTATIONAL SETUP section). It is clear from these 

results that the interaction tones make the dominant contribution 

to the overall tone noise signature of the open rotor for all the 

speeds considered in this study. 
NASA/TM—2014-218132 14 

A. D. 

B. E. 

C. F. 

Figure 20. Variations of predicted overall sound pressure level (OASPL) as a function of the sideline angle 

for all aft rotor tones (i.e., ∑ nBPF2) and interaction tones (i.e., ∑ mBPF1+kBPF2) considered in this study. 

Results are shown for highest tip speed considered (i.e., 6436 RPMC) to the lowest (4620 RPMC). 
NASA/TM—2014-218132 15 

DISCUSSION

In view of the acoustic results shown in Figures 16 through 

20, it is the main conclusion of this study that any open rotor 

noise reduction strategy that ignores the importance of the 

interaction tones may not only be sub-optimal, it may actually 

lead to inadvertent noise level increases. That is because, while 

the rotor tones and their harmonics are controlled by the 

outboard portion of the span, the interaction tones receive far 

more distributed contributions from the blade span. In fact, the 

distribution is not even uniform in that loss of a portion of the 

blade could lead to higher noise levels at some sideline angles. 

Therefore, noise control strategies involving the clipping of the 

aft blade tip or modifying the front rotor blade tips to reduce the 

strength of the interaction of the front blade tip vortex with the 

aft blade tips are only effective in reducing the nBPF2 tone 

levels. Such approaches, however, would do little to reduce the 

interaction tones, which are controlled by the wake of the front 

rotor. Therefore, to substantially reduce open rotor tone noise 

beyond what has been achieved through blade tip clipping or 

similar approaches requires complementary techniques 

involving judicious control of the wake of the front rotor. An 

example of such an approach is the so-called trailing edge 

blowing (15). Furthermore, wake management strategies should 

be predicated on optimization studies that maximize noise 

reductions for the target tones while minimizing noise increases 

for the other tones. It should be noted that spacing increases 

could also reduce wake strength, but it is likely that the 

necessary spacing increases would be impractical. 

CONCLUSIONS 

Three-dimensional PIV measurements have been used to 

validate FINETM/Turbo CFD simulations of a contra-rotating 

open rotor system. Detailed comparisons of the measured and 

predicted velocity fields show good agreement for the evolution 

of the forward rotor wake/tip vortex characteristics. Surface 

pressure distributions from the CFD simulations were used as 

input to the LINPROP acoustic code to predict the associated 

open rotor tone noise spectra. 

While the absolute noise level predictions have proven to 

be difficult, the sideline directivity trends were shown to be 

quite reasonable. These results provided confidence in the 

predictions so that they can be used to gain better insight into 

the radiated fields by the various open rotor tones. Insights that 

are not available from the measured data. In particular, the 

results from the LINPROP prediction code indicate that, in 

contrast to the individual rotor tone levels which are controlled 

by the blade outboard portion, the interaction tones levels are 

controlled by the entire span suggesting that blade wakes are 

more important than the tip vortices in setting the interaction 

tone levels. This insight in turn suggests that a comprehensive 

approach to open rotor noise reduction should involve not only 

tip vortex interaction mitigation, but also blade wake mitigation 

to enable further noise reduction for open rotor systems. 

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