David M. Elliott 

Glenn Research Center, Cleveland, Ohio

Initial Investigation of the Acoustics of a CounterRotating Open Rotor Model With Historical

Baseline Blades in a Low-Speed Wind Tunnel

NASA/TM—2012-217258

March 2012

AIAA–2011–2760
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David M. Elliott 

Glenn Research Center, Cleveland, Ohio

Initial Investigation of the Acoustics of a CounterRotating Open Rotor Model With Historical

Baseline Blades in a Low-Speed Wind Tunnel

NASA/TM—2012-217258

March 2012

AIAA–2011–2760

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

17th Aeroacoustics Conference

cosponsored by the American Institute of Aeronautics and Astronautics

and the Confederation of European Aerospace Societies

Portland, Oregon, June 5–8, 2011
Acknowledgments

The NASA Environmentally Responsible Aviation Project provided funding for testing support while the

NASA Subsonic Fixed Wing Project provided funding for analysis support.

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either expressed or implied, by the National Aeronautics and 

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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—2012-217258 1 

Initial Investigation of the Acoustics of a CounterRotating Open Rotor Model With Historical Baseline 

Blades in a Low-Speed Wind Tunnel 

David M. Elliott 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

A counter-rotating open rotor scale model was tested in the NASA Glenn Research Center 9- by 

15-Foot Low-Speed Wind Tunnel (LSWT). This model used a historical baseline blade set with which 

modern blade designs will be compared against on an acoustic and aerodynamic performance basis. 

Different blade pitch angles simulating approach and takeoff conditions were tested, along with angle-ofattack configurations. A configuration was also tested in order to determine the acoustic effects of a 

pylon. The shaft speed was varied for each configuration in order to get data over a range of operability. 

The freestream Mach number was also varied for some configurations. Sideline acoustic data were taken 

for each of these test configurations. 

Nomenclature 

ORPR Open Rotor Propulsion Rig 

OAPWL Overall Power Level 

BPF Blade Passage Frequency 

OASPL Overall Sound Pressure Level 

SPL Sound Pressure Level 

PSD Power Spectral Density 

LSWT Low Speed Wind Tunnel 

AOA Angle of Attack 

Introduction 

Counter-rotating open rotor models were tested at the NASA Glenn Research Center in the 1980’s 

and early 1990’s (Refs. 1 and 2). At that time, the fuel consumption advantage of open rotors was not 

great enough to overcome their other shortcomings compared to ducted fans. One of the primary 

shortcomings was noise output. Due to the increasing cost of fuel in recent years, there is a renewed 

interest in open rotor propulsion systems. Fortunately, there are new improved technologies and tools 

available, which can be applied to the design of open rotors, which will help mitigate their disadvantages. 

NASA is again investigating the counter-rotating open rotor to determine if it is a viable propulsion 

system for future aircraft. In partnership with General Electric Aviation, the NASA Glenn Research 

Center started a test program to test advanced blade sets on the scale Open Rotor Propulsion Rig (ORPR). 

This rig was tested in the NASA Glenn Research Center 9- by 15-Foot Low-Speed Wind Tunnel (LSWT) 

in order to get acoustic and aerodynamic performance data. This paper focuses on the acoustic data 

acquired during the NASA portion of the test (Ref. 3), which was funded by the Environmentally 

Responsible Aviation Project. The main goal of this part of the overall test campaign was to acquire data 

on a so-called historical baseline blade set. The historical baseline blade set is representative of late 

1980’s state of the art, when open rotor systems were last tested on a wide scale. This blade set 

concentrated on aerodynamic performance as opposed to acoustic noise reduction. Advanced blade sets 
NASA/TM—2012-217258 2 

will be compared against the historical baseline blades to determine if there are any acoustic or 

aerodynamic improvements. The historical baseline blade data is also used to validate prediction models 

for counter-rotation open rotor simulations. 

Test Hardware 

The ORPR was the same drive system that had been used in previous testing. It was refurbished for 

this and future tests. The ORPR model is of the pusher type wherein the rotors are located toward the 

back of the turbomachinery. Turbines within the drive rig are fed by a 450 lb/in.2

 air source, which turns 

the rotor blades. The drive rig is also mounted on a turntable, which provides angle of attack (AOA) 

testing capability. The capability to simulate a pylon was also designed and installed in some 

configurations. The blades in the historical baseline set, also designated F31/A31, are approximately 

25.7 in. in diameter for the forward rotor and 24.8 in. in diameter for the aft rotor. The forward rotor has 

12 blades and the aft rotor has 10 blades. The pitch of the blades could be adjusted to provide different 

operational conditions. Two pitch angle configurations were tested during this entry corresponding to 

takeoff and approach power conditions. The takeoff configuration had pitch angles of 40.1 for the 

forward rotor and 40.8 for the aft rotor. The approach configuration had pitch angles of 33.5 for the 

forward rotor and 35.7 for the aft rotor. The pitch angles were measured at a radius of 9.6 in. These pitch 

angles were chosen in order to give the best aerodynamic performance at a given shaft speed and flight 

condition of interest while allowing the rotor torque split to be equal. For this entry, all test points 

matched forward and aft rotor shaft speed. A photograph of the blades is shown in Figure 1. The 

installation of the model and drive rig in the 9- by 15-Foot Low-Speed Wind Tunnel is shown in Figures 

2 and 3. Figure 3 also shows the installed pylon hardware. 

Test Matrix 

The test matrix reported here consisted of four separate configurations. These configurations included 

two runs at takeoff blade pitch angle and two runs at approach blade pitch angle. Each blade pitch angle 

was run with and without the pylon. For a subset of speeds, data for additional angles of attack of 3 and 

8 were obtained. The majority of the test runs were at a wind tunnel freestream Mach number of 0.2, 

however for three of the four runs data was also obtained at the tunnel Mach number of 0.22. Data were 

acquired at corrected model speeds from 70.7 to 114.4 percent of design. This corresponds to corrected 

shaft speeds of 4628 to 7487 rpm. The test matrix is summarized in Table 1. During this portion of the 

test, rotors were run at equal speeds. 

Test Facility 

This test was conducted in the NASA Glenn Research Center 9- by 15-Foot Low-Speed Wind Tunnel 

(Refs. 4 to 6). The 9- by 15-Foot Low-Speed Wind Tunnel has acoustically treated walls using Kevlar 

batting. The maximum flow velocity is approximately 0.22 to 0.23 Mach. 

Sideline Acoustics 

Sideline acoustic measurements were obtained using a traversing microphone probe on a track 

parallel to the fan axis and 60 in. away from it. Data were taken at 18 positions on the traverse at 

7.5 intervals ranging from 17.8 to 140.6 relative to the rear rotor pitch change axis with zero reference 

upstream. The microphone used was 1/4-in. in diameter and a nose cone was installed for in-flow 

measurements. Data were taken at a sampling rate of 200 kHz in order to ensure a frequency upper limit 

of at least 60 kHz. This would permit scaling this data to full-scale for system analysis purposes. 

Corrections to the data have been made for microphone response, bullet nose receptivity, atmospheric 
NASA/TM—2012-217258 3 

attenuation, and spherical spreading. All far-field data shown are corrected to 1-ft lossless distance. The 

accuracy of the data system is 1 dB. The traversing microphone probe installed in the 9- by 15-Foot 

Low-Speed Wind Tunnel is shown in Figure 3. The 18 traversing microphone stops and corresponding 

measured angles relative to the rear rotor pitch change axis are listed in Table 2. Note that the traversing 

microphone starts in the aft position and moves forward during a test. 

Results 

Sideline Acoustic Results 

Representative Spectra 

It is useful to look at a sample spectrum to get an idea of what the noise signature of the counter-rotating 

open rotor is. Figure 4 shows the spectrum of the historical baseline blades at 6450 corrected rpm using the 

takeoff pitch setting and with the microphone at 141 relative to the rear pitch change axis. The large tonal 

content can be seen from 0 to 20 kHz with the majority in the 0 to 10 kHz range. This is much different 

from counter-rotating open rotors with equal blade count. From 20 kHz and higher, broadband noise is 

dominant but drops off steadily. This shows that it is important to focus on the noise below 20 kHz as far as 

tonal content is concerned. Figure 5 shows the same condition as Figure 4, but now the frequency range has 

been narrowed from 0 to 10 kHz in order to better illustrate the significant tonal content. Tones are present 

over this entire frequency range. Figure 6 shows this same spectra zoomed in further from 0 to 5 kHz. Here 

the blade passage and interaction tones are individually labeled. The subscript letter “f” after BPF is for the 

forward rotor and likewise the “a” is for the aft rotor. While the fundamental front and aft rotor blade 

passage tones are relatively high, on the order of 20 dB above broadband, their higher harmonics are much 

lower, less than 10 dB above broadband. The highest tone present is the interaction tone BPFf + 2BPFa

which is approximately 25 dB higher than broadband level. The other interaction tones 2BPFf + BPFa and 

BPFf + 3BPFa are as high as BPFa. BPFf + BPFa and 2BPFf + 2BPFa are at least 15 dB higher than 

broadband levels. This illustrates the interaction tone dominance that is a feature of open rotors with nonequal blade counts at approach and takeoff conditions. Since one-third octave filters are commonly used in 

acoustic data reduction, a one third-octave band spectrum from 0 to 20 kHz of the data in Figure 4 is shown 

in Figure 7. This one-third octave spectrum focuses on the fact that the majority of the tonal energy is 

contained within the bands that make up the 2000 to 4000 Hz range. 

Directivity 

Figure 8 shows noise spectra at directivity angles of 45, 90, and 141 for the takeoff pitch angle and 

6450 corrected rpm. In terms of broadband content, the 90 spectrum is the lowest while the 141

spectrum is on the order of 2 dB higher over most of the frequency range. The 45 spectrum is similar in 

broadband to the 141 spectrum until about 4 kHz where it drops to 90 spectrum level. When 

considering tones, it is difficult to glean a definitive trend due to the significant variability of tone levels 

in the spectra for the three different angles. Figure 9 shows the data for the approach pitch angle and 5598 

corrected rpm. At this forward angle, once again the 90 spectrum shows the lowest broadband, but now 

the 45 spectrum is higher than the 141 spectrum for frequencies up to 10 kHz where the two spectral 

levels converge. Below 10 kHz the difference between broadband levels is on the order of 2 dB. The 141

spectrum appears to have greater tonal content than the spectra at the other two angles for this lower 

speed condition. While individual spectra can be a useful tool for examining detailed noise, it is beneficial 

to look at the directivity of overall sound pressure level. These data are shown in Figure 10 for the takeoff 

and approach pitch angles. When calculating OASPL, the frequency range from 500 Hz to 50 kHz was 

used in order to include the lowest blade passage tones while staying above the wind tunnel’s anechoic 

lower limit. It can be seen in Figure 10 that the takeoff pitch angle OASPL is, on average, 15 dB higher 

than that for the approach pitch. The directivity shape for the approach pitch angle shows that the OASPL 
NASA/TM—2012-217258 4 

is higher at the lowest and highest directivity angles and dipping in the middle. In fact, the highest 

OASPL is at the forward most angles for approach blade pitch, approximately 8 dB as compared to broad 

angles. The OASPL for takeoff pitch angle shows much less of a variance between upstream and 

downstream angles as compared to broadside angles, only 2 to 3 dB. The inset of Figure 10 shows a plan 

view of the ORPR with the reference of the directivity angles. Figure 11 shows the directivity for the 

highest interaction tone seen in Figure 6, the BPFf + 2BPFa tone. For this tone there is a much higher 

variability at takeoff than approach. The variability at takeoff is approximately 10 dB while approach is 

only 5 dB. 

Shaft Speed and Blade Pitch Angle 

Examining how the open rotor blades behave at different shaft speeds is useful since this will show 

noise output over an operating speed line. The shaft speed is also related to blade pitch angle since blade 

pitch needs to be optimally set for a given shaft speed in order to provide maximum aerodynamic 

performance. Due to time limitations, it was not possible to obtain more than the two blade pitch angles 

shown in the test matrix, but these two pitch angles can be compared over the speed range for which data 

were obtained for them. Figure 12 shows overall acoustic power levels as a function of rpm for the 

takeoff and approach pitch angles. The design points are shown at the upper end of the takeoff pitch curve 

and the lower end of the approach pitch curve. The level difference between these conditions is 

approximately 16 dB. While the curves have very similar slopes, since the approach curve lies to the right 

of the takeoff, it has a lower overall acoustic power level for any given shaft speed. 

The spectra in Figures 13 and 14 show the progression up the speed line for the takeoff pitch angle for 

the 141 and 45 directivity angles, respectively. At 141, increasing speed approximately 1000 rpm 

raises broadband noise level by about 10 dB. This is also true at 45 for the higher speeds, but somewhat 

less for the lower speed change, on the order of 7 dB. It is also interesting to examine how the pitch angle 

affects the spectrum at approximately the same shaft speed. The closest shaft speed match for the takeoff 

and approach pitch angles is 6316 and 6325 rpm respectively. Figures 15 and 16 show the spectra 

comparisons for different pitch angles at these shaft speeds for 141 and 45. At 141 directivity, the 

broadband is higher for the takeoff pitch by about 13 dB. It is interesting to note that the tones are 

somewhat higher relative to broadband levels for the approach pitch angle. The same results appear at the 

45 angle although the amount of tonal content is somewhat less for the approach pitch angle. 

Angle of Attack 

The variance in angle of attack was also examined. Figure 17 shows how the overall power levels 

change with speed for three angles of attack at the takeoff blade pitch angle. There is a small difference 

between 0 and 3, particularly close to the design condition at higher speeds. The 8 angle of attack level 

is 2 to 3 dB higher over the entire speed range. Figure 18 shows this same type of data for the approach 

pitch angle. Once again, the level for the 3 angle of attack does not vary much compared with the 0

level while the 8 level is 2 dB higher over the speed lines. Example spectra for the angle of attack 

variations at takeoff blade pitch angle are shown in Figures 19 and 20 for 141 and 45 directivity, 

respectively. These spectra are for the corrected shaft speed of 6316 rpm, since this speed was one of the 

few cases for which data was acquired for all three angles of attack. The increase in broadband noise is 

clearly seen as angle of attack is increased. It appears that there is more tonal content at 8 angle of attack 

relative to the 0 and 3 cases. 
NASA/TM—2012-217258 5 

Freestream Mach Number 

Some limited data were obtained for a freestream wind tunnel Mach number of 0.22 as opposed to 

0.2. This 0.22 Mach number was closer to actual flight conditions. As seen in the test matrix speeds, the 

corrected shaft speed was adjusted to correspond to the higher Mach number condition. Figure 21 

compares overall sound power levels for the speed lines of the two different Mach numbers using the 

takeoff pitch angle. At the lower speed where the rotors are off condition, freestream Mach number 

conditions have the same overall power levels. At higher speeds where the rotors are closer to design 

condition, the higher freestream Mach number of 0.22 produces lower power levels by about 1.5 dB. 

Figure 22, which plots the same data for the approach pitch angle, shows the higher Mach number of 0.22 

produces overall power levels which are lower by approximately 1 dB. The spectral plots of Figures 23 

and 24 show the difference in freestream Mach numbers for the 141 and 45 directivity angles using the 

approach pitch angle, where the larger difference was seen in the power levels. The shaft speeds are 

6773 rpm for Mach 0.2 and 6617 rpm for Mach 0.22. The 141 spectrum shows the lower broadband 

level occurring at 5000 Hz and up. In contrast, there is little difference for the 45 spectrum. These results 

suggest that most of the reduction in the higher Mach number occurs in the aft directivity area. 

Pylon Effects 

Some very important data were acquired with a simulated pylon installed upstream of the rotor 

blades. It is important to determine what effects the pylon has on the extra noise produced due to the 

ingestion of the pylon wakes by the rotors, especially the front rotor. This should better represent the 

noise generated by the counter rotating open rotor when in an actual installed pusher configuration. 

Figure 25 compares overall acoustic power levels as a function of shaft speed at the takeoff pitch angle 

for isolated and pylon configurations. At the higher speeds, including those close to design condition, the 

pylon introduces a low level of extra noise. For example, at the highest speed this is only 0.5 dB. At 

further off design conditions corresponding to lower speeds, the pylon produces higher power levels, on 

the order of 3 dB. Figure 26 compares these power levels for the approach pitch angle. Here the overall 

trend is similar in that the pylon produces less noise as the shaft speed approaches the design value. 

However in this instance the pylon produces more noise levels at and close to design condition compared 

with the takeoff condition. In fact the power levels produced by the pylon are 3 to 4 dB higher in the low 

speed conditions. Flow field measurements may help determine why the lower speeds are more affected 

by pylon noise. These measurements could show if the pylon wakes decay faster or are more modest at 

higher speeds. These results are once again examined by looking at spectra. Figures 27 and 28 show the 

141 and 45 spectra at the takeoff pitch angle and 6450 rpm speed for isolated and pylon configurations. 

As expected, there is not a great deal of difference in these spectral levels since this is a high-speed 

condition. The pylon does appear to produce more tonal content at 45 in the 1 to 4 kHz range, however, 

other frequency ranges show little tonal differences. The broadband levels are also very similar. 

Figures 29 and 30 show the 141 and 45 spectra at the approach pitch angle and 5598 rpm speed for 

isolated and pylon configurations. As is expected, there are larger differences seen in these spectra as 

compared to the takeoff condition. Lower broadband levels are evident at both 141 and 45 and this 

difference is on the order of 1 to 2 dB. Tonal content also appears to be slightly greater for the pylon 

configuration for both angles. Clearly, more research is warranted to better understand the effect of the 

pylon for various blade settings and power conditions. 

Conclusion 

The extensive acoustic data obtained for the historical baseline open rotor blade set can be used to 

benchmark future advanced blade designs on the basis of noise and aerodynamic performance. The 

historical blade set, owing to its non-equal blade counts, indicates the dominance of interaction tones 

relative to the individual rotor tones. In general, the historical baseline blade set tends to radiate more 
NASA/TM—2012-217258 6 

overall acoustic power levels to the upstream and downstream angles compared with the broad side 

angles. This noise directivity feature is more pronounced at the approach pitch angle. Noise power levels 

also increase steadily as shaft speed is increased. Shaft speed and blade pitch angles are interrelated and 

the takeoff pitch angle has higher noise levels for a given rpm than the approach pitch angle. Angle of 

attack also affects the noise output. Angles of attack up to 3 has little effect on the noise power levels, 

while an angle of attack of 8 produces a substantial increase on the order of 2 to 3 dB. Testing at a 

freestream Mach number of 0.22 instead of 0.2, and adjusting rpm accordingly shows a reduction of the 

noise power levels by approximately 1 dB. Spectra tend to show that most of the benefit of the higher 

freestream Mach number occurs in the aft directivity angles. The addition of a pylon increases the noise 

level and this is more so at lower speeds than higher speeds. The increase in noise due to the pylon at the 

takeoff condition is less than 1 dB. However at the approach condition the pylon noise increase is as 

much as 4 dB. Other measurement techniques such as flow field particle imaging velocimetry, when used 

in conjunction with this acoustic data, could help explain the causes of some of the phenomenon observed 

in this test. Future blade designs can utilize the results of this data by concentrating efforts to reduce noise 

of open rotors. 

Figure 1.—The historical baseline blades used in the open rotor propulsion rig. 
NASA/TM—2012-217258 7 

Figure 2.—The open rotor propulsion rig and traversing microphone 

probe installed in the NASA Glenn 9- by 15-Foot Low-Speed Wind 

Tunnel. 

Figure 3.—The open rotor propulsion rig installed in the NASA Glenn 

9- by 15-Foot Low-Speed Wind Tunnel with pylon hardware 

mounted. 
NASA/TM—2012-217258 8 

TABLE 1.—THE TEST MATRIX OF TEST SPEEDS RUN FOR THE OPEN ROTOR NASA ENVIRONMENTALLY 

RESPONSIBLE AVIATION PROJECT IN THE NASA GLENN 9- BY 15-FOOT LOW-SPEED WIND TUNNEL 

Corrected

Speeds

Tested

(RPM)

Takeoff Pitch –

Pylon

Approach Pitch –

Pylon

Approach Pitch –

Isolated

Takeoff Pitch ‐ 

Isolated

Mach 0.2

AOA 0

4628, 5561, 6080,

6316, 6450

5598, 6325, 6773,

7245, 7487

5598, 6325, 6773,

7245, 7487

4628, 5277, 5561,

6080, 6316, 6450

Mach 0.2

AOA 3

4628, 5561, 6316 5598, 6325, 6773 5598, 6325, 6773 4628, 5561, 6316

Mach 0.2

AOA 8

4628, 5561 5598, 6325, 6773 5598, 6325, 6773 4628, 5561, 6316

Mach 0.22

AOA 0

5903, 6617, 7054 5903, 6617, 7054 4880, 5790, 6530

TABLE 2.—MEASURED GEOMETRIC ANGLES RELATIVE TO THE REAR ROTOR 

PITCH CHANGE AXIS FOR EACH MICROPHONE TRAVERSING PROBE STOP 

Traversing microphone stop Angle, 

degrees 

18 17.8 

17 22.5 

16 30 

15 37.5 

14 45 

13 52.5 

12 60 

11 67.5 

10 75 

9 82.5 

8 90 

7 97.5 

6 105 

5 112.5 

4 120 

3 127.5 

2 135 

1 140.6 
NASA/TM—2012-217258 9 

Figure 4.—Power spectrum density of the counter-rotating open 

rotor historical baseline blades at 6450 corrected rpm with 

takeoff pitch angle and 141 relative to the rear rotor pitch 

change axis. 

Figure 5.—Power spectrum density from 0 to 10 kHz of the counterrotating open rotor historical baseline blades at 6450 corrected 

rpm with takeoff pitch angle and 141 relative to the rear rotor 

pitch change axis. 
NASA/TM—2012-217258 10 

Figure 6.—Power spectrum density from 0 to 5 kHz of the counter-rotating open rotor historical baseline 

blades at 6450 corrected rpm with takeoff pitch angle and 141 relative to the rear rotor pitch change axis. 

blade passage and interaction tones are labeled. 

Figure 7.—One-third octave sound pressure level of the counterrotating open rotor historical baseline blades at 6450 corrected rpm 

with takeoff pitch angle and 141 relative to the rear rotor pitch 

change axis. 
NASA/TM—2012-217258 11 

Figure 8.—Power spectrum density showing three different directivity 

angles relative to the rear rotor pitch change axis for takeoff pitch 

angle and 6450 corrected rpm. 

Figure 9.—Power spectrum density showing three different directivity 

angles relative to the rear rotor pitch change axis for approach pitch 

angle and 5598 corrected rpm. 
NASA/TM—2012-217258 12 

Figure 10.—Overall sound pressure level (500 Hz to 50 kHz) directivity 

for approach pitch angle at 5598 corrected rpm and takeoff pitch angle 

at 6450 corrected rpm. The reference of directivity to the ORPR model 

is shown in the inset. 

Figure 11.—Power spectrum density directivity of the forward blade 

passage + 2(Aft blade passage) interaction tone for approach 

pitch angle at 5598 corrected rpm and takeoff pitch angle at 6450 

corrected rpm. 
NASA/TM—2012-217258 13 

Figure 12.—Overall power levels (500 Hz to 50 kHz) versus corrected rpm 

of the historical baseline blades for takeoff and approach pitch angles. 

The 98.5 and 29 percent design speeds are also shown. 

Figure 13.—Power spectrum density for three corrected rpm’s of 

the historical baseline blades at takeoff pitch angle and 141

relative to the rear rotor pitch change axis. 
NASA/TM—2012-217258 14 

Figure 14.—Power spectrum density for three corrected rpm’s of 

the historical baseline blades at takeoff pitch angle and 45

relative to the rear rotor pitch change axis. 

Figure 15.—Power spectrum density comparison of takeoff and 

approach pitch angles for similar corrected rpm conditions at 

141 relative to the rear rotor pitch change axis. 
NASA/TM—2012-217258 15 

Figure 16.—Power spectrum density comparison of takeoff and 

approach pitch angles for similar corrected rpm conditions at 45

relative to the rear rotor pitch change axis. 

Figure 17.—Overall power levels (500 Hz to 50 kHz) versus 

corrected rpm for three angles of attack of the historical baseline 

blades with takeoff pitch angle. 
NASA/TM—2012-217258 16 

Figure 18.—Overall power levels (500 Hz to 50 kHz) versus corrected 

rpm for three angles of attack of the historical baseline blades with 

approach pitch angle. 

Figure 19.—Power spectrum density of the historical baseline blades 

for three angles of attack at takeoff pitch angle, 6316 corrected 

rpm, and 141 relative to the rear rotor pitch change axis. 
NASA/TM—2012-217258 17 

Figure 20.—Power spectrum density of the historical baseline blades 

for three angles of attack at takeoff pitch angle, 6316 corrected 

rpm, and 45 relative to the rear rotor pitch change axis. 

Figure 21.—Overall Power Level (500 Hz to 50 kHz) versus 

corrected rpm of Historical Baseline Blades for two free flow 

Mach numbers and takeoff pitch angle. The rpm was adjusted 

for 0.22 Mach to match condition. 
NASA/TM—2012-217258 18 

Figure 22.—Overall power level (500 Hz to 50 kHz) versus corrected 

rpm of historical baseline blades for two free flow Mach numbers 

and approach pitch angle. The rpm was adjusted for 0.22 Mach to 

match condition. 

Figure 23.—Power spectrum density of historical baseline blades for 

two free flow Mach numbers at approach pitch angle and 141

relative to rear rotor pitch change axis. The rpm was 6773 for 

Mach 0.2 and 6617 for Mach 0.22. 
NASA/TM—2012-217258 19 

Figure 24.—Power spectrum density of historical baseline blades 

for two free flow Mach numbers at approach pitch angle and 45

relative to rear rotor pitch change axis. The rpm was 6773 for 

Mach 0.2 and 6617 for Mach 0.22. 

Figure 25.—Overall power level (500 Hz to 50 kHz) versus 

corrected rpm of historical baseline blades for isolated and pylon 

configurations at takeoff pitch angle. 
NASA/TM—2012-217258 20 

Figure 26.—Overall power level (500 Hz to 50 kHz) versus corrected 

rpm of historical baseline blades for isolated and pylon configurations 

at approach pitch angle. 

Figure 27.—Power spectrum density of historical baseline blades 

for isolated and pylon configurations at takeoff pitch angle, 6450 

corrected rpm, and 141 relative to rear rotor pitch change axis. 
NASA/TM—2012-217258 21 

Figure 28.—Power spectrum density of historical baseline blades for 

isolated and pylon configurations at takeoff pitch angle, 6450 corrected 

rpm, and 45 relative to rear rotor pitch change axis. 

Figure 29.—Power spectrum density of historical baseline blades for 

isolated and pylon configurations at approach pitch angle, 5598 

corrected rpm, and 141 relative to rear rotor pitch change axis. 
NASA/TM—2012-217258 22 

Figure 30.—Power spectrum density of historical baseline blades for 

isolated and pylon configurations at approach pitch angle, 5598 

corrected rpm, and 45 relative to rear rotor pitch change axis. 

References 

1. Woodward, R.P., “Noise of a Model High Speed Counter-rotation Propeller at Simulated 

Takeoff/Approach Conditions (F7/A7),” NASA TM-100206, Oct. 1987. 

2. Woodward, R.P., Hall, D.G., Podboy, G.G., and Jeracki, R.J., “Takeoff/Approach Noise for a Model 

Counter-rotation Propeller with a Forward-Swept Upstream Rotor,” AIAA-93-0596, presented at the 

31st Aerospace Sciences Meeting & Exhibit, Reno, NV, January 11–14, 1993. 

3. Van Zante, Dale E., Gazzaniga John A., Elliott, David M., and Woodward, Richard P., “An Open 

Rotor Test Case: F31/A31 Historical Baseline Blade Set,” ISABE-2011-1310, 2011, to be published.

4. Dahl, M.D., and Woodward, R.P., “Comparison Between Design and Installed Acoustic 

Characteristics of the NASA Lewis 9-by 15-Foot Low-Speed Wind Tunnel Acoustic Treatment,” 

NASA TP-2996, Apr. 1990.

5. Dahl, M.D., and Woodward, R.P., “Background Noise Levels Measured In the NASA Lewis 9- by 

15-Foot Low-Speed Wind Tunnel,” NASA TP-3274, Nov. 1992. 

6. Woodward, R.P., and Dittmar, J.H., “Background Noise Levels Measured in the NASA Lewis 9- by 

15- Foot Low-Speed Wind Tunnel,” NASA TM-106817/AIAA-95-0720, 1995. 
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01-03-2012 

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Technical Memorandum 

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4. TITLE AND SUBTITLE

Initial Investigation of the Acoustics of a Counter-Rotating Open Rotor Model With 

Historical Baseline Blades in a Low-Speed Wind Tunnel 

5a. CONTRACT NUMBER

5b. GRANT NUMBER

5c. PROGRAM ELEMENT NUMBER

6. AUTHOR(S)

Elliott, David, M. 

5d. PROJECT NUMBER

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5f. WORK UNIT NUMBER

WBS 699959.02.09.03.05 

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 

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E-18013 

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National Aeronautics and Space Administration 

Washington, DC 20546-0001 

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NASA 

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NASA/TM-2012-217258 

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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

A counter-rotating open rotor scale model was tested in the NASA Glenn Research Center 9- by 15-Foot Low-Speed Wind Tunnel (LSWT). 

This model used a historical baseline blade set with which modern blade designs will be compared against on an acoustic and aerodynamic 

performance basis. Different blade pitch angles simulating approach and takeoff conditions were tested, along with angle-of-attack 

configurations. A configuration was also tested in order to determine the acoustic effects of a pylon. The shaft speed was varied for each 

configuration in order to get data over a range of operability. The freestream Mach number was also varied for some configurations. 

Sideline acoustic data were taken for each of these test configurations.

15. SUBJECT TERMS

Acoustics; Counter rotating; Open rotor blades; Baseline blades 

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