David B. Stephens

Glenn Research Center, Cleveland, Ohio

Nearfi eld Unsteady Pressures at Cruise Mach 

Numbers for a Model Scale Counter-Rotation 

Open Rotor

NASA/TM—2012-217672

September 2012

AIAA–2012–2264
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
David B. Stephens

Glenn Research Center, Cleveland, Ohio

Nearfi eld Unsteady Pressures at Cruise Mach 

Numbers for a Model Scale Counter-Rotation 

Open Rotor

NASA/TM—2012-217672

September 2012

AIAA–2012–2264

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

18th Aeroacoustics Conference

cosponsored by the American Institute of Aeronautics and Astronautics and Confederation of 

European Aerospace Societies

Colorado Springs, Colorado, June 4–6, 2012
Acknowledgments

This report was funded by the NASA Environmentally Responsible Aviation (ERA) project and 

the NASA Subsonic Fixed Wing (SFW) project.

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

Level of Review: This material has been technically reviewed by technical management. 
Nearfield Unsteady Pressures at Cruise Mach Numbers

for a Model Scale Counter-Rotation Open Rotor

Abstract

An open rotor experiment was conducted at cruise Mach numbers and the unsteady pressure in

the nearfield was measured. The system included extensive performance measurements, which can help

provide insight into the noise generating mechanisms in the absence of flow measurements. A set of

data acquired at a constant blade pitch angle but various rotor speeds was examined. The tone levels

generated by the front and rear rotor were found to be nearly equal when the thrust was evenly balanced

between rotors.

Nomenclature

f Frequency, Hz

B Number of rotor blades

c Speed of sound, m/s

D Rotor diameter, m

J Rotor advance ratio, V0/NπD

M Mach number, V /c

m Front rotor blade rate harmonic number

N Rotor shaft rotation frequency, Hz

n Aft rotor blade rate harmonic number

SO Shaft Order, f /N

T Thrust, N

Q Torque, Nm

V Velocity, m/s

η Propeller efficiency, T V0/QN

θ Sideline angle measured from upstream along the model centerline

Subscripts

a Aft rotor

e Emitted

f Forward rotor

g Geometric

0 Freestream conditions

1 Introduction

The nearfield unsteady pressure field near a model scale open rotor was evaluated in the 8x6 Supersonic Wind

Tunnel at NASA Glenn Research Center (GRC). High-speed propellers and open rotors offer considerable

fuel efficiency benefits compared with turbofans, but these systems present noise challenges to both the

community around the airport and to persons on the airplane. A recent test campaign was conducted as a

collaboration between NASA and General Electric Aviation (GE) to document the performance and acoustics

of open rotors. Advanced blade designs have been developed by GE since the previous test campaign in the

1980’s. The primary incentive for implementing open rotor propulsion is the fuel efficiency at cruise Mach

numbers, and this tunnel was operated between Mach 0.27 and Mach 0.85. During the test, a streamwise

NASA/TM—2012-217672 1

David B. Stephens 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 
line array of unsteady pressure transducers was mounted to a plate above the model which could be lowered

from the tunnel ceiling toward the blade tip. These data could be used for estimating structural loading on

the airframe or for estimating en route noise, for example.

Open rotors were previously studied in the 8x6 Supersonic Wind Tunnel during the mid 1980’s. The

results from those tests were documented in a number of reports including those by Dittmar[1], Dittmar and

Stang[2], Dittmar et al.[3][2] and Whitfield et al.[4].

2 Experiment

The tests described in this report were conducted in the 8x6 wind tunnel from February to September 2011.

As described below, these tests included scaled aerodynamic performance, structurally responsive blades,

and realistic acoustic signatures. Comparisons of model scale acoustic data to the full-scale counter rotating

engine flight test in the 1980’s were found to be in good agreement.[5]

2.1 Drive Rig

Figure 1: Open Rotor Propulsion Rig and F31/A31

blade set in 8x6 Wind Tunnel at NASA Glenn Research

Center. NASA Image C-2010-3454.

The NASA counter rotating open rotor propulsion rig (ORPR) was used to power the open rotor

blades in this experiment. The rig consists of two

counter rotating spools with a non-rotating center

shaft. Each of the two counter rotating spools is

attached to a two-stage air turbine on the aft end

that can produce up to 560 kW (750 HP). The

two air turbines are supplied by 2 MPa (300 psi)

air heated to 90◦C (200◦F), and can be independently controlled using a feed-back system to drive

the rotor speeds independently to more than 8,000

RPM with an accuracy of ± 15 RPM. This drive

rig was operated in the late 1980’s, and was refurbished in order to correct any detrimental effects of

long-term storage. The refurbishment activity consisted of a general inspection and rebuild of the mechanical components, construction of a new swept

pylon, refurbishment of the forward and aft rotating force balances and installation of a telemetry

system. This rig was operated in a low-speed test

campaign from late 2009 to late 2010 in the 9x15

Low Speed Wind Tunnel at NASA Glenn. Reports

to data on data from the low speed portion of this

project include Elliott[6] and Stephens and Envia[7]

and Berton[8]. The refurbished ORPR installed in

the 8x6 Wind Tunnel is shown in Figure 1.

2.2 Historical Baseline Blades (F31 / A31)

The model test hardware simulates a counter rotating open rotor engine design. New hubs were manufactured

as part of the current test campaign, with support for 12 blades for the forward rotor and 10 blades for the

aft rotor. The pitch angle setting for both blade rows are manually adjustable. The blade set used for the

tests described in the present paper was designed and manufactured by GE and is known as F31/A31. This

blade set represents a good early 1990’s aerodynamic design, but is not acoustically optimized. The F31/A31

blade set was manufactured with modern materials and techniques for this test, and will be used to provide

a baseline to anchor new data. The forward rotor is 65.18 cm (25.66 inches) in diameter and the aft rotor is

62.97 cm (24.79 inches) in diameter. Axial spacing between the two rotors is variable, but was set at 19.91

cm (7.84 inches) between the pitch change axis of each rotor for the data presented in this report. The center

NASA/TM—2012-217672 2
of the aft rotor at the blade pitch change axis location is the origin of the coordinate system for the results

shown in this paper, except where noted.

2.3 8x6 Wind Tunnel

The 8x6 Wind Tunnel at GRC is a high speed wind tunnel with velocity capabilities from Mach 0.27 to

Mach 2.0. The test section is 2.44 m high by 1.83 m wide (8 feet by 6 feet) and 7.16 m (23.5 feet) long. The

test section has perforated bleed holes for boundary layer removal that are used when the tunnel is operated

at transonic velocities. The flow is driven by three electric motors that can produce up to 65 MW (87,000

HP). A description of the 8x6 Wind Tunnel complex is given by Soeder[9].

2.4 Acoustic Instrumentation Table 1: Location of unsteady

pressure transducers relative to

aft rotor pitch change axis.

Sensor Axial Location

Number cm (in)

1 -46.6 (-18.36)

2 -39.0 (-15.36)

3 -34.1 (-13.44)

4 -29.6 (-11.64)

5 -23.5 (-9.24)

6 -18.9 (-7.44)

7 -14.8 (-5.84)

8 -7.0 (-2.76)

9 0 (0)

10 7.0 (2.76)

11 14.8 (5.84)

12 18.9 (7.44)

13 23.5 (9.24)

14 29.6 (11.64)

15 34.1 (13.44)

16 39.0 (15.36)

17 46.6 (18.36)

The unsteady pressure instrumentation was an aluminum plate with 17

flush-mounted XCS-093-15SG transducers made by Kulite Semiconductor. The transducers offer a 15 PSI range and are temperature compensated. The calibration values provided with the instrumentation were

used to convert the voltage output to unsteady pressure. These sensors

were connected to a Precision Filter and recorded by a RC Electronics

DataMAX. The pressure transducers were simultaneously sampled at 200

kHz in 15 second long records. These records were processed into narrowband SPL using the Digital Acoustics Data Software (DADS) package

developed at GRC. The instrumented plate could be remotely lowered

from the tunnel ceiling to achieve various sideline distances from the

model axis of rotation. A table of the geometry is in Table 1. These

locations are relative to the location of the aft rotor pitch change axis.

A schematic of the instrumentation relative to the model turboprop is

shown in Figure 2.

2.5 Operating Conditions

An extensive range of operating conditions were tested, with blade pitch

angles for the forward and aft blades ranging from 43.0◦/43.5◦ at Mach

0.27-0.45 to 64.4◦/61.8◦ at Mach 0.73-0.85. Powered rotation speeds

ranged up to 8250 RPM. The thrust and torque of each rotor were measured by custom built rotating balances mounted in the hub of each rotor.

Maximum thrust was limited by the instrumentation to 2 kN (450 lbf) per rotor while torque was limited

to 680 Nm (500 ft-lb) per rotor.

A large number of conditions were tested, with several hundred readings of data taken with the F31/A31

blade set. Most of these were taken with the Kulite plate in the stowed position against the test section

ceiling, so are not particularly interesting for unsteady pressure measurements, due primarily to the limited

directivity that could be measured in this configuration. At Mach 0.4 with the plate in the stowed position,

the sensors only span emitted angles between 46 and 90 degrees. With the plate lowered to 0.78Df above

the model centerline, at Mach 0.78 the directivity spans emitted angles of 12 to 97 degrees from upstream.

The rest of the data presented in this report is from the model at these operating conditions.

A set of readings was chosen as representing a particularly interesting subset of the data. These readings

represent a series of increasing rotor speeds at Mach 0.78 that include a local peak in the propulsive efficiency.

The ORPR does not have a powered or automated mechanism for changing the rotor blade pitch during a

run, so only a limited number of iterations on blade pitch was practical during the course of testing. Changes

to the rotor RPM at a fixed tunnel velocity V0 provide a substitute method for changing the blade inflow

angles. A few performance metrics for these readings are given in Table 2, where the total thrust is given as a

fraction of the maximum total rating for the thrust balances (4 kN). Details of the performance calculations

are given in the report by Hughes and Gazzaniga.[10] While it would be desirable from a scientific standpoint

NASA/TM—2012-217672 3
Figure 2: Illustration of the unsteady pressure instrumentation in the 8x6 wind tunnel. Plate is shown at

the sideline distance of 0.78Df . Measurement locations are numbered from downstream.

Table 2: Open Rotor Performance at Mach 0.78. βf = 64.4, βa = 61.8, Hplate = 0.78Df .

Escort RDG RPMf RPMa (Tf + Ta)/Tmax Tf /Ta Qf /Qa Jf η

2938 5714 5713 0.1 8.14 2.85 1.38 0.600

2943 6094 6093 0.32 1.23 1.35 1.29 0.803

2948 6336 6337 0.48 0.96 1.11 1.25 0.828

2953 6475 6474 0.58 0.89 1.04 1.22 0.830

2958 6664 6664 0.71 0.83 0.98 1.18 0.826

2966 6943 6943 0.89 0.80 0.95 1.14 0.812

to evaluate a complicated system like the ORPR with each parameter varied individually, the cost of testing

and the large number of parameters combined to prevent this sort of detailed study.

3 Unsteady Pressure Results and Discussion

The 8x6 test section is made of stainless steel plate perforated with angled holes for boundary layer bleed

during transonic operation. Needless to say, it is not an ideal acoustic environment. The redeeming aspects

of this test are the high Mach number flow which pushes reflected sound waves downstream, and the excellent

signal-to-noise ratio available due to the placement of the sensors in the immediate vicinity of the model.

Equation 1 can be used for calculating the frequency of interaction tones expected by an open rotor

system,

Shaft Order Interaction Tones = mBfNf + nBaNa

.5 (Nf + Na)

, (1)

where m and n are integers denoting the front and rear blade passing tone order respectively. This formula

provides a good approximation to the tone order when the rotational speeds of the two rotors are nearly the

same.

3.1 Background Noise Removal

An extensive database of “tare run” measurements was acquired. These are measurements made with the

rig in the test section, but without blades installed or compressed air delivered to the turbine. Blade fillers

are used to fill the gaps in the hub in the absence of rotor blades.

A typical narrowband (12.2 Hz) pressure spectrum from sensor 9 is presented in Figure 3. Most of the

strongest tones are at multiples of the aft rotor blade passing frequency. The plate is at location 0.78Df

NASA/TM—2012-217672 4
from the center of the rig. The plot shows both a tare run spectrum and a powered spectrum for Mach 0.78.

The tones generated by the open rotor are immediately apparent, and no signal to noise problem is evident

for these spectral features. The broadband signal generated by the open rotor is clearly above the tare run

spectra for frequencies above 3 kHz for the rotor speed shown.

0 10 20 30 40 50 60 70 80

110

115

120

125

130

135

140

145

150

155

Shaft Order

SPL, dB

Powered

Tare

Aft BPF

Foreward BPF

Figure 3: Nearfield unsteady pressure spectra at 90 degrees. M=0.78. Powered data at 6943 RPM.

An additional complication exists because the data were not acquired at constant emission angles or

constant geometric angles, but rather at constant sideline locations. Additionally, the 17 sensors make for

a somewhat sparse data set. This makes it challenging to compare measurements made a different sideline

distances, since the geometric and emitted angles both change with sideline measurement location. The

relationship between geometric and emission angles was calculated using,

θe = θg − arcsin (M0 sin (θg)), (2)

which assumes linear propagation.

3.2 Measured Pressure Spectra

As can be seen in Figure 3, the magnitude of the tare run spectra may be higher than that of the powered

spectra for some frequencies. A background separation routine was developed, and subtraction at each

frequency was carried out. If the magnitude of the appropriate tare run spectrum was less than 6 dB below

the magnitude of the powered spectrum, the resulting value was set to not-a-number (NAN) representing an

ambiguous value. These NAN values are not shown in the remaining acoustic results.

Figure 4 shows the integrated unsteady pressure generated on a sideline by the rotor at Mach 0.78 for

the six conditions listed in Table 2. For these results, the tare run spectra were subtracted, a value of 0 dB

was substituted for the NAN values and the results were integrated from 500 to 50,000 Hz. The aft rotor is

centered on x = 0 by definition and the forward rotor is centered at x = 0.2 m. Note that the direction of the

x-axis has been reversed for Figure 4 to correspond better with the model orientation and sensor numbering

shown in Figure 2. For each value of Jf shown in Figure 4, two peaks can be observed with a local minimum

in between. For the the two highest advance ratios (lowest RPM) the forward peak is higher, while the aft

peak is higher otherwise. This corresponds reasonably with the torque and thrust ratios shown in Table 2.

The frequency content of each peak in the OASPL was considered in more detail. These peaks correspond

to sensors 9 and 11, as described in Table 1. The narrowband sound spectra were processed using a peakfinding algorithm to identify tones which were at least 6 dB above the lowest magnitude level around every

integer shaft order. If a tone was found, the spectral energy in that tone was summed. The tone content was

NASA/TM—2012-217672 5
Table 3: Decibel levels of blade rate tones for forward and aft rotor.

Advance Ratio Jf 1.38 1.29 1.25 1.22 1.18 1.14

Sensor 11, SO = 12 145.8 151.3 156.1 157.5 159.2 162.1

Sensor 9, SO = 10 143.0 150.8 157.5 159.2 159.7 162.9

usually spread over three or four frequency bins, due to slight drifting of the rotor speed. If a peak was not

found, an average broadband level was calculated. If only values of NAN were found, indicating levels below

the tare run level, NAN was returned. This data was used to create the bar charts of Figure 5 and Figure 6,

which denote the dB level (ref 20μ Pa) of the RMS pressure at integer shaft orders. This processing method

was found to be useful for quantifying both tone and broadband levels in one plot, while still accounting for

NAN values in the spectra.

Aft rotor tones dominate the spectra measured by sensor 9, as shown in Figure 5, while interaction tones

(mBf +nBa) are a much smaller contribution. The second aft tone harmonic was an average of 6.8 dB below

the first for the six conditions considered, while the third tone dropped another 4.7 dB on average. The

forward rotor tones were at least 9.1 dB below the aft tone of the same order, for the first three harmonics.

Forward rotor tones dominate the spectra measured by sensor 11, as shown in Figure 6. Interaction tones

again were not found to be a significant contribution. The forward tone harmonics were found to decrease

differently than the aft tones, as the second forward tone harmonic was an average of 4.6 dB below the first

and the third tone dropped another 7.5 dB on average. The aft rotor tones were a minimum of 5.5 dB below

the forward tone of the same order, for the first three harmonics.

The directivity of the first forward and aft blade passing tones were also examined. These tones were

chosen for closer examination because they are among the strongest and because these lower frequencies

have larger wavelengths and so are better resolved by the sparse spatial measurement. A quick consideration

of the emitted angles revealed the need to consider the forward and aft tones as being emitted from two

different spatial locations. The close proximity of the measurement sideline makes this significant. For an

observer much further away, the emitted angles would tend towards the same values for each rotor. In Figure

7, the emitted angles for the forward and aft rotor tones are plotted with respect to the center of the forward

and aft rotors, respectively. On these figures, the center sensor corresponds to the data presented in Figure

5 and is called out with a large marker. This geometry definition causes the radiation patters to collapse

quite nicely, as shown in Figure 7 (a), (c) and (e). No magnitude correction was made for these results,

as the measurements were clearly made in the acoustic near-field and using free-field distance corrections

would not be meaningful. The corresponding mean squared coherence is shown in (b), (d) and (f) of Figure

7for the same three values of advance ratio. The coherence was calculated by using the sensor closest to

40 degrees emission angle as the reference, with the appropriate origin for each rotor tone as previously

described. This emission angle generally measured the highest levels for these two tones. At the highest

RPM (lowest advance ratio for constant Mach number) conditions, these tones are coherent over a very wide

angle. Abrupt changes in the coherence are likely due to reflections of pressure waves between the aluminum

plate and the ORPR.

The peak levels for the two blade rate tones were found to be remarkably close in magnitude, although

measured by different sensors. These values are given in Table 3. Aside from Jf = 1.38, which corresponds

to a thrust ratio of Tf /Ta = 8.14, these tone levels are within 1.7 dB. The diminutive contribution of the

interaction tones is likely due to the smaller diameter of the aft rotor, which is designed to reduce the

interaction of the aft rotor with the blade tip vortices from the front rotor. This effect was noted by Dittmar

and Stang[2] and by Woodward and Gordon[11].

4 Conclusions

An extensive test campaign considering a model scale open rotor at cruise conditions has been completed,

and a brief overview of the unsteady pressure results has been provided. The OASPL on a 0.78Df sideline

exceeds 175 dB, with the blade rate tone and harmonics responsible for the majority of this signal. These

pressure spectra have high coherence (MSC > .9) over emitted angles between 25 and 75 degrees.

NASA/TM—2012-217672 6
0.4 0.3 0.2 0.1 0 −0.1 −0.2 −0.3 −0.4

140

145

150

155

160

165

170

175

180

Axial Location, m

OASPL, dB

Upstream Downstream

(a) Jf = 1.38

0.4 0.3 0.2 0.1 0 −0.1 −0.2 −0.3 −0.4

140

145

150

155

160

165

170

175

180

Sideline Distance, m

OASPL, dB

(b) Jf = 1.29

0.4 0.3 0.2 0.1 0 −0.1 −0.2 −0.3 −0.4

140

145

150

155

160

165

170

175

180

Sideline Distance, m

OASPL, dB

(c) Jf = 1.25

0.4 0.3 0.2 0.1 0 −0.1 −0.2 −0.3 −0.4

140

145

150

155

160

165

170

175

180

Sideline Distance, m

OASPL, dB

(d) Jf = 1.22

0.4 0.3 0.2 0.1 0 −0.1 −0.2 −0.3 −0.4

140

145

150

155

160

165

170

175

180

Sideline Distance, m

OASPL, dB

(e) Jf = 1.18

0.4 0.3 0.2 0.1 0 −0.1 −0.2 −0.3 −0.4

140

145

150

155

160

165

170

175

180

Sideline Distance, m

OASPL, dB

(f) Jf = 1.14

Figure 4: OASPL levels along the measurement sideline for each sensor at the six conditions shown in Table

2. Note that the direction of the x-axis has been reversed to correspond better with the model orientation

and sensor numbering shown in Figure 2

NASA/TM—2012-217672 7
0 10 20 30 40 50

120

125

130

135

140

145

150

155

160

165

Shaft Order

SPL, dB

Broadband

Aft Rotor Tones

Front Rotor Tones

Interaction Tones

(a) Jf = 1.38

0 10 20 30 40 50

120

125

130

135

140

145

150

155

160

165

Shaft Order

SPL, dB

(b) Jf = 1.29

0 10 20 30 40 50

120

125

130

135

140

145

150

155

160

165

Shaft Order

SPL, dB

(c) Jf = 1.25

0 10 20 30 40 50

120

125

130

135

140

145

150

155

160

165

Shaft Order

SPL, dB

(d) Jf = 1.22

0 10 20 30 40 50

120

125

130

135

140

145

150

155

160

165

Shaft Order

SPL, dB

(e) Jf = 1.18

0 10 20 30 40 50

120

125

130

135

140

145

150

155

160

165

Shaft Order

SPL, dB

(f) Jf = 1.14

Figure 5: Spectral content of unsteady pressure measured by sensor number 9 at various advance ratios,

Mach 0.78. θg = 90 degrees. θe = 39 degrees.

NASA/TM—2012-217672 8
0 10 20 30 40 50

120

125

130

135

140

145

150

155

160

165

Shaft Order

SPL, dB

Broadband

Aft Rotor Tones

Front Rotor Tones

Interaction Tones

(a) Jf = 1.38

0 10 20 30 40 50

120

125

130

135

140

145

150

155

160

165

Shaft Order

SPL, dB

(b) Jf = 1.29

0 10 20 30 40 50

120

125

130

135

140

145

150

155

160

165

Shaft Order

SPL, dB

(c) Jf = 1.25

0 10 20 30 40 50

120

125

130

135

140

145

150

155

160

165

Shaft Order

SPL, dB

(d) Jf = 1.22

0 10 20 30 40 50

120

125

130

135

140

145

150

155

160

165

Shaft Order

SPL, dB

(e) Jf = 1.18

0 10 20 30 40 50

120

125

130

135

140

145

150

155

160

165

Shaft Order

SPL, dB

(f) Jf = 1.14

Figure 6: Spectral content of unsteady pressure measured by sensor number 11 at various advance ratios,

Mach 0.78. θe = 74 degrees. θe = 25 degrees.

NASA/TM—2012-217672 9
20 40 60 80 100 120

120

125

130

135

140

145

150

155

160

165

Emitted Angle, degrees

Tone Level, dB

Aft Rotor BPF, SO = 10

Front Rotor BPF, SO = 12

(a) Jf = 1.29

20 40 60 80 100 120

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1

Emitted Angle, degrees

Mean Squared Coherence

(b) Jf = 1.29

20 40 60 80 100 120

120

125

130

135

140

145

150

155

160

165

Emitted Angle, degrees

Tone Level, dB

(c) Jf = 1.22

20 40 60 80 100 120

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1

Emitted Angle, degrees

Mean Squared Coherence

(d) Jf = 1.22

20 40 60 80 100 120

120

125

130

135

140

145

150

155

160

165

Emitted Angle, degrees

Tone Level, dB

(e) Jf = 1.14

20 40 60 80 100 120

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

1

Emitted Angle, degrees

Mean Squared Coherence

(f) Jf = 1.14

Figure 7: Directivity and coherence of aft and forward blade passing frequency tones. Emission angle has

been calculated from the center of the respective rotor. Large symbol indicates center sensor. M=0.78.

NASA/TM—2012-217672 10
References

[1] Dittmar, J. H., “Cruise Noise of Counterrotation Propeller at Angle of Attack in Wind Tunnel,”

NASA/TM-88869 , October 1986.

[2] Dittmar, J. H. and Stang, D. B., “Noise Reduction for Model Counterrotation Porpeller at Cruise by

Reducing Aft-Propeller Diameter,” NASA TM-88936 , May 1987.

[3] Dittmar, J. H., Gordon, E. B., and Jeracki, R. J., “The Effect of Front-to-Rear Propeller Spacing on

the Interaction Noise at Cruise Conditions of a Model Counterrotation Propeller Having a Reduced

Diameter Aft Propeller,” NASA TM-101329 , October 1988.

[4] Whitfield, C. E., Mani, R., and Gliebe, P. R., “High Speed Turboprop Aeroacoustic Study (Counterrotation) Volume I - Model Development,” NASA CR-185241 , July 1990.

[5] Woodward, R. P., Loeffler, I. J., and Dittmar, J. H., “Measured Far-Field Flight Noise of a Counterrotation Turboprop at Cruise Conditions,” NASA/TM-101383 , January 1989.

[6] Elliott, D. M., “Initial Investigation of the Acoustics of a Counter Rotating Open Rotor Model With

Historical Baseline Blades in a Low Speed Wind Tunnel. AIAA-2011-2760,” 17th AIAA/CEAS Aeroacoustics Conference,, Portland, Oregon, USA, 5-8 June 2011.

[7] Stephens, D. B. and Envia, E., “Acoustic Shielding for a Model Scale Counter-rotation Open Rotor.

AIAA-2011-2940,” 17th AIAA/CEAS Aeroacoustics Conference,, Portland, Oregon, USA, 5-8 June

2011.

[8] Berton, J. J., “Empennage Noise Shielding Benefits for an Open Rotor Transport. AIAA-2011-2764,”

17th AIAA/CEAS Aeroacoustics Conference,, Portland, Oregon, USA, 5-8 June 2011.

[9] Soeder, R. H., “NASA Lewis 8- by 6-Foot Low-Speed Wind Tunnel User Maunal,” NASA/TM 105771 ,

February 1993.

[10] Hughes, C. E. and Gazzaniga, J. A., “Summary of Low-Speed Wind Tunnel Results of Several HighSpeed Counterrotation Propeller Configurations. AIAA-1988-3149,” AIAA/ASME/SAE/ASEE 24th

Joint Propulsion Conference,, Boston, Massachusetts, USA, 11-13 July 1988.

[11] Woodward, R. P. and Gordon, E. B., “Noise of a Model Counterrotation Propeller With Reduced Aft

Rotor Diameter at Simulated Takeoff/Approach Conditions. AIAA-1988-0263,” AIAA 26th Aerospace

Sciences Meeting,, Reno, Nevada, USA, 11-14 January 1988.

NASA/TM—2012-217672 11
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)

01-09-2012 

2. REPORT TYPE

Technical Memorandum 

3. DATES COVERED (From - To)

4. TITLE AND SUBTITLE

Nearfield Unsteady Pressures at Cruise Mach Numbers for a Model Scale Counter-Rotation 

Open Rotor 

5a. CONTRACT NUMBER

5b. GRANT NUMBER

5c. PROGRAM ELEMENT NUMBER

6. AUTHOR(S)

Stephens, David, B. 

5d. PROJECT NUMBER

5e. TASK NUMBER

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 

8. PERFORMING ORGANIZATION

 REPORT NUMBER

E-18347 

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-2012-217672 

12. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified-Unlimited 

Subject Categories: 71 and 01 

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

An open rotor experiment was conducted at cruise Mach numbers and the unsteady pressure in the nearfield was measured. The system 

included extensive performance measurements, which can help provide insight into the noise generating mechanisms in the absence of flow 

measurements. A set of data acquired at a constant blade pitch angle but various rotor speeds was examined. The tone levels generated by 

the front and rear rotor were found to be nearly equal when the thrust was evenly balanced between rotors. 

15. SUBJECT TERMS

Counter rotation; Rotors; Wind tunnel tests 

16. SECURITY CLASSIFICATION OF: 17. LIMITATION OF

 ABSTRACT

UU 

18. NUMBER

 OF 

 PAGES

18 

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