Richard F. Bozak and Christopher E. Hughes

Glenn Research Center, Cleveland, Ohio

James Buckley

Vantage Partners, Brook Park, Ohio 

The Aerodynamic Performance of an 

Over-the-Rotor Liner With Circumferential 

Grooves on a High Bypass Ratio Turbofan Rotor

NASA/TM—2013-218066

August 2013

GT2013–95114
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Richard F. Bozak and Christopher E. Hughes

Glenn Research Center, Cleveland, Ohio

James Buckley

Vantage Partners, Brook Park, Ohio 

The Aerodynamic Performance of an 

Over-the-Rotor Liner With Circumferential 

Grooves on a High Bypass Ratio Turbofan Rotor

NASA/TM—2013-218066

August 2013

GT2013–95114

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

Turbo Expo 2013

sponsored by the American Society of Mechanical Engineers (ASME)

San Antonio, Texas, June 3–7, 2013
Acknowledgments

The authors would like to thank Daniel Tweedt of AP Solutions for his aerodynamic analysis and design of the over-the-rotor rub 

strip. Thanks also to the W-8 test cell crew: Julia Stephens, Brent Seifert, Bruce Wright, James Knight, and Jason Bryant. This 

work was supported by the Environmentally Responsible Aviation Project of NASA’s Integrated Systems Research Program.

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

Trade names and trademarks are used in this report for identification 

only. Their usage does not constitute an official endorsement, 

either expressed or implied, by the National Aeronautics and 

Space Administration.

Level of Review: This material has been technically reviewed by technical management. 
NASA/TM—2013-218066 1

The Aerodynamic Performance of an Over-the-Rotor Liner With 

Circumferential Grooves on a High Bypass Ratio Turbofan Rotor

Richard F. Bozak and Christopher E. Hughes

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

James Buckley

Vantage Partners

Brook Park, Ohio 44142

Abstract

While liners have been utilized throughout turbofan ducts to 

attenuate fan noise, additional attenuation is obtainable by 

placing an acoustic liner over-the-rotor. Previous experiments 

have shown significant fan performance losses when acoustic 

liners are installed over-the-rotor. The fan blades induce an 

oscillating flow in the acoustic liners which results in a 

performance loss near the blade tip. An over-the-rotor liner 

was designed with circumferential grooves between the fan 

blade tips and the acoustic liner to reduce the oscillating flow 

in the acoustic liner. An experiment was conducted in the W-8 

Single-Stage Axial Compressor Facility at NASA Glenn 

Research Center on a 1.5 pressure ratio fan to evaluate the 

impact of this over-the-rotor treatment design on fan 

aerodynamic performance. The addition of a circumferentially

grooved over-the-rotor design between the fan blades and the 

acoustic liner reduced the performance loss, in terms of fan 

adiabatic efficiency, to less than 1 percent which is within the 

repeatability of this experiment.

Introduction

While the use of air transportation continues to grow, the 

reduction of negative environmental impacts will become 

increasingly challenging. The Environmentally Responsible 

Aviation (ERA) Project of NASA’s Integrated Systems 

Research Program is working to develop technologies to 

reduce the impact of commercial aircraft on the environment. 

Specifically, the project is working to improve fuel efficiency, 

reduce emissions, and reduce noise. The noise from turbofan 

engines on commercial aircraft is a major contributor to 

airport community noise. Increasing bypass ratios have 

elevated the impact of fan noise relative to the other noise 

sources in turbofan engines. Acoustic liners have been 

developed and utilized to reduce both tonal and broadband 

noise levels within the fan duct.

While liners within fan ducts have successfully reduced 

engine noise, additional acoustic benefits are possible if a liner 

is installed over-the-rotor. A series of acoustics and fan 

performance tests at NASA Glenn Research Center have 

explored the impact of installing a foam-metal liner over-therotor. First, a foam-metal liner was developed for broadband 

attenuation and tested on a low-speed fan. This liner reduced

fan noise by 4 dB in broadband sound pressure level (Ref. 1).

Then in 2008, an over-the-rotor foam-metal liner was installed 

on an ultra-high bypass fan model in the 9x15 Low-Speed 

Wind Tunnel. The results from this test showed a reduction of

up to 1 dB in overall acoustic power level (OAPWL) (Ref. 2), 

with a loss in fan aerodynamic efficiency between 3.75 and 

8.75 percent (Ref. 3). In addition to the significant loss in 

efficiency, the perforated plate covering the foam-metal liner 

caused damage to the composite fan blade tips. Another foammetal liner over-the-rotor was evaluated on a static engine test 

of a Williams International FJ44-3A (Ref. 4). The foam-metal 

liner was installed over-the-rotor and behind holes in the 

valleys of the circumferential grooves. The static engine test 

showed an OAPWL reduction of 2.5 dB, while reducing 

engine performance between 0.5 and 2 percent. The ability of 

this treatment design to reduce noise, while having less of an 

effect on fan performance, when compared to other over-therotor liners, has been attributed to the addition of

circumferential grooves around the fan casing.

Circumferentially grooved casing treatments were examined 

in the 1970s because of their ability to increase the flow range 

capability of single-stage axial compressors with no 

significant effect on efficiency (Ref. 5). Increasing the flow 

range enables fan or compressor operation at higher speed 

conditions with sufficient stall margin, which could result in a 

net increase in efficiency. Hathaway (Ref. 6) summarized the 

important over-the-rotor circumferential groove parameters. In 

order for a design to improve the flow range without 

significantly impacting efficiency, the middle 60 percent of 

the rotor chord should be treated, the grooves should have 65

to 75 percent open area, and the grooves should be 3 times as 

deep as they are wide (Ref. 6). The ability of circumferential 

grooves to improve the flow range without significantly 

impacting efficiency has led to its use in modern turbofans.

An over-the-rotor treatment was designed for testing in the 

9x15 Low-Speed Wind Tunnel. This treatment incorporates

circumferential grooves over-the-rotor. The over-the-rotor 

grooves were designed and modeled by Daniel Tweedt of AP 
NASA/TM—2013-218066 2

Solutions Inc. In order to reduce the risk of finding significant 

performance losses in an expensive wind tunnel test, the overthe-rotor treatment design was tested in the W-8 Single Stage 

Axial Compressor Facility. The test included pressure and 

temperature measurements to examine the performance losses.

The results of this test are presented.

Nomenclature

SDT Source Diagnostic Test

OAPWL overall acoustic power level

RPMc corrected revolutions per minute

PR total pressure ratio

FPR area averaged fan pressure ratio

TR total temperature ratio

FTR area averaged fan temperature ratio

γ ratio of specific heats

ηadiabatic adiabatic efficiency

FAE area averaged fan adiabatic efficiency

Test Facility

The test was conducted in the W-8 Single-Stage Axial 

Compressor Facility at NASA Glenn Research Center. A 

schematic of the facility is shown in Figure 1. The facility is 

capable of delivering up to 7,000 hp (5,220 kW) at speeds up 

to 21,240 rpm to a 22 in. (0.559 m) diameter fan or

compressor. Up to 100 lbm/s (45.4 kg/s) of air is provided from 

an atmospheric inlet, or a dry pressurized air source. Flow 

conditioning screens in the inlet plenum reduce the turbulence 

intensity at the fan to less than 1 percent. In order to provide a 

flight-like boundary layer entering the fan, up to 4.5 lbm/s 

(2.0 kg/s) of air is pulled through a bleed slot about 10 in.

(0.25 m) upstream of the fan. The flow through the facility is 

controlled by a sleeve throttle valve. The air can be exhausted 

through an atmospheric exhaust system or an altitude exhaust 

system. Facility instrumentation includes up to 400 channels 

of steady pressure and thermocouple measurement, tip 

clearance sensing, and up to 96 channels of high speed 

rotating fan data. The steady pressure and temperature 

measurements presented were acquired at 1 Hz and then 

averaged over a 10 seconds. 

Since the primary objective of this test was to determine the 

performance loss associated with a new over-the-rotor 

treatment design, a subset of the facility capabilities was 

utilized. In order to increase the facility flow capacity, the 

inlet orifice plate was removed. The boundary layer bleed was 

used to full capacity, in order to mimic the conditions in the 

9x15 Low-Speed Wind Tunnel. The corrected mass flow 

removed through the bleed slot was approximately 4.5 lbc/s

(2.0 kgc/s) for all conditions presented. Only the atmospheric 

inlet and atmospheric exhaust were used for this test.

Figure 1.—Schematic of the W-8 Single Stage Axial Compressor Facility at NASA Glenn Research Center.
NASA/TM—2013-218066 3

Test Hardware

The Source Diagnostic Test (SDT) fan rotor (Refs. 7 and 8)

was used for this test. The fan, designated R4, was designed by 

General Electric Aircraft Engines and is a 1/5 scale model of a 

high bypass turbofan. The fan has a 22-in. (0.559 m) diameter 

with 22 titanium blades, a hub to tip ratio of 0.3, and a design 

speed of 12,657 RPMc. The fan has a nominal hot tip clearance 

of 0.020 in. (0.51 mm) and tip speed of 1215 ft/s (370.3 m/s) at 

the design speed. This fan was chosen for its titanium blades, 

since over-the-rotor treatments have been shown to damage 

composite blades (Ref. 2). The data from previous testing 

(Refs. 7 and 8) was used to avoid stall so that blade stress

measurements would not be needed during this test.

The test hardware consisted of two configurations as shown 

in Figure 2. The baseline was tested first and then last to check 

repeatability. This configuration contained a smooth, 

aluminum rub strip without grooves or acoustic treatment. The 

only over-the-rotor instrumentation in this configuration was a

set of four blade tip proximity sensors equally spaced around 

the circumference and on the blade stacking line.

The fan case containing the acoustic treatment contains two 

treated sections; one over-the-rotor and one upstream of the 

fan. Circumferential grooves cover the entire blade chord

over-the-rotor. The grooves have 67 percent open area and 

the groove height increases from about 2 times the groove 

width at the blade leading edge to about 3 times the groove 

width at the trailing edge. Slots in the bottoms of the grooves 

allow acoustic waves to enter the acoustic treatment.

A region upstream of the fan contained another set of 

similarly sized chambers arranged in 12 rows. These chambers 

were open directly to the flow surface, behind a layer of fibermetal, unprotected by grooves.

Instrumentation

Since the facility orifice plate was removed to increase the 

flow capacity, an alternative mass flow measurement was used. 

The mass flow was then measured using a correlation of 

corrected mass flow to the inlet static pressure (Ref. 5). This 

correlation has a total uncertainty of 1.0 percent. The mass flow 

removed through the boundary layer bleed slot is measured with 

an orifice plate. The bleed slot flow subtracted from the inlet 

mass flow calculation gives the mass flow through the fan.

Blade tip clearance sensors were installed to center the 

baseline case. Two thermocouples were placed in the treatment 

as a precautionary measure, to prevent the plastic stereolithic 

treatment from reaching its glass transition temperature. These 

thermocouples were placed at two axial locations; one at the fan 

face and one near the fan trailing edge.

Figure 2.—Cross section of the fan with the over-the-rotor hardware identified.
NASA/TM—2013-218066 4

Figure 3.—Radial traversing total pressure, total 

temperature probe.

Static and total pressure measurements were made with the 

Electronically Scanned Pressure (ESP) system. These pressure 

measurements have an uncertainty of ±0.005 psi (34.5 Pa). 

Static pressure ports were located on the fan case as well as on 

the hub. The inlet total pressure and total temperature were 

measured at four circumferential stations in the plenum. The 

average of these four measurements is the inlet total pressure 

and total temperature used to calculate the fan’s pressure and 

temperature rise. Temperatures were measured with Type E 

thermocouples with an uncertainty of ±2 °R (1.1 °C). Fan 

performance measurements were taken approximately 4.5 in.

(11 cm) downstream the fan face. The fan performance 

measurement plane contains three total pressure rakes, three

total temperature rakes, and a single radially traversing 

pressure and temperature probe. The pressure and temperature 

rakes each contain seven radially spaced, equal area ports. The 

traversing probe incorporates a single total pressure and total 

temperature measurement, as shown in Figure 3. A schematic 

of the fan duct with instrumentation locations is shown in

Figure 4.

Figure 4.—Cross-section of the fan duct showing instrumentation locations.
NASA/TM—2013-218066 5

Results

The objective of this test was to quantify the difference 

between a baseline hard-wall rub strip and a rub strip with an 

over-the-rotor treatment. A set of fan and flow conditions 

were chosen for comparison with historical data. The 

difference in fan pressure ratio, fan temperature ratio, and fan 

adiabatic efficiency are presented to quantify the performance 

loss associated with the use of the over-the-rotor treatment.

The fan case was designed with a split section over-therotor to simplify the configuration change between the 

baseline and treated case. By avoiding the need to remove 

instrumentation between configurations, the repeatability of 

the results was improved. In order to quantify the repeatability

of the measurements presented, the test of the baseline case

was conducted before and then repeated after the test of the 

treatment case. The repeatability data is presented to verify the

performance differences.

The conditions were set along the nominal fixed nozzle 

operating line for the SDT fan defined in References 7 and 8). 

First, the corrected fan speed was set. Then, the sleeve throttle 

valve, located downstream of the test section, was adjusted until 

the corrected mass flow was obtained. The corrected mass flow 

was set to within 0.3 lbc/s (0.14 kgc/s) of their target value. 

These corrected fan speed conditions are shown in Table 1. 

Conditions between 50 and 87.5 percent design speed were used 

to represent the tip speed and fan pressure ratio of an ultra-high 

bypass turbofan. The 61.7 and 87.5 percent design speed were 

used to compare to the approach and cutback data from 

previous testing (Refs. 7 and 8).

The ratio of fan exit total pressure to the inlet total pressure is 

the pressure ratio, while the ratio of fan exit total temperature to 

the inlet total temperature is the temperature ratio. The fan 

pressure ratio and fan temperature ratio are the area averaged 

pressure and temperature ratios, respectively. The absolute 

uncertainties associated with each calculation are shown in 

Table 1. In order to estimate the precision/random error, 27 

repeated measurements at three conditions were used to 

compute a 95 percent confidence interval for each parameter. 

These measurements were back-to-back repeats of the pressure 

and temperature rake data obtained during consecutive 

traversing probe measurements. While the absolute 

uncertainties for the measurements are large, the repeatability 

suggests that comparisons between the two rub strip 

configurations with the same instrumentation in the same 

locations are accurate within 1 percent of each measurement.

Figure 5 shows the radial variation in total pressure and 

total temperature ratio as measured from the average of three

rakes in (a) and (b) or from a 28 point total pressure and total 

temperature probe traverse in (c) and (d). The results of both 

the rake averaged and detailed probe total pressure 

measurements, Figure 5(a) and (c), show that there is not 

significant change in total pressure ratio due to the addition of 

the over-the-rotor treatment. These results are consistent with 

the results obtained by Sutliff, et al., (Ref. 4) where a similar 

over-the-rotor design was used. However, the averaged total 

temperature rake measurements, Figure 5(b), show a 

temperature increase near the blade tips when the over-therotor treatment is installed, while from mid-span to the hub the 

temperature difference is insignificant. The total temperature

probe measurements in Figure 5(d) show that the temperature 

increase extends from the tip to about 88 percent span at 

77.5 percent of design speed.

The adiabatic efficiency (ηadiabatic) at each measurement 

location is calculated from the pressure and temperature ratios 

by the formula shown below, where γ is the ratio of specific 

heats for air. 

( )

( 1)

1

adiabatic − η =

γ γ−

TR

PR

(1)

The fan adiabatic efficiency (FAE) is then computed as the 

area averaged adiabatic efficiency from the total pressure and 

total temperature rake measurements. The computed FAEs 

over the range of conditions in Table 1 are shown in Figure 6. 

The FAE is compared against data taken with the same fan 

and duct geometry from the 9x15 Low-Speed Wind Tunnel in 

2002 (Ref. 6). The test of the baseline case test repeated within 

0.005 up to 77.5 percent of design speed. This repeatability 

suggests that the actual uncertainty might be less than that 

calculated in Table 1, but more testing would be needed to 

verify. As the fan speed is increased, the loss in efficiency for 

the over-the-rotor treatment case reduces from 0.02 at 

50 percent of design speed to 0.0075 at 87.5 percent of design 

speed. However, all of the differences in efficiency are within 

the uncertainty shown in Table 1.

TABLE 1.—AEROPERFORMANCE ERRORS AND UNCERTAINTIES AT FIXED NOZZLE,

NOMINAL OPERATING LINE FLOW CONDITIONS

% design 

speed

RPMc Fan total pressure ratio (PR) 

relative errors

Fan total temperature ratio (TR) 

relative errors

Fan adiabatic efficiency (FAE) 

relative errors

Absolute uncertainty, 

%

Repeatability, 

%

Absolute uncertainty, 

%

Repeatability, 

%

Absolute uncertainty, 

%

Repeatability, 

%

61.7 7809 ±0.52 ±0.04 ±16.0 ±0.03 ±13.9 ±0.77

77.5 9809 ±0.31 ±0.06 ±10.7 ±0.04 ±8.9 ±0.65

87.5 11075 ±0.23 ±0.11 ±8.4 ±0.07 ±7.2 ±0.60
NASA/TM—2013-218066 6

(a) (b)

(c) (d)

Figure 5.—Fan exit pressure and temperature profiles; rake averaged total pressure ratio (a) and total temperature ratio (b), 

and traversing probe total pressure ratio (c) and total temperature ratio (d).
NASA/TM—2013-218066 7

Figure 6.—Comparison of fan adiabatic efficiency on 

the nominal fixed nozzle operating line.

Figure 7.—Comparison of fan exit and over-the-rotor 

treatment temperatures with increasing circumferential 

blade tip Mach number.

Thermocouples in the acoustic treatment showed a 

significant increase in temperature when the blade tip speed 

approached the speed of sound. The increase in total 

temperature ratio of the treatment relative to that of the rake 

measurements is shown in Figure 7. The rapid increase of 

treatment temperature prevented the acquisition of 

performance data above 87.5 percent of design speed for the 

treated fan case. While the fan exit rake temperatures and 

treatment temperature at the fan face steadily increase, the 

treatment temperature near the fan trailing edge dramatically 

increases when the circumferential blade tip Mach number is 

above 0.8. 

Conclusion

An experiment was conducted in the W-8 Single Stage 

Axial Compressor Facility at NASA Glenn Research center to 

evaluate the fan performance loss of an over-the-rotor liner 

with a circumferentially grooved design between the blade tips 

and the acoustic treatment. The results of the experiment were 

repeatable to within 0.6 percent of fan adiabatic efficiency and 

consistent with the results from previous testing of the same 

1.5 pressure ratio fan. While the fan pressure ratio was not 

affected by the over-the-rotor treatment, an increase in fan 

temperature ratio was seen near the blade tips. The fan 

performance loss was shown to be as low as 0.75 percent in 

terms of fan adiabatic efficiency. While the loss in adiabatic 

efficiency is within the absolute measurement uncertainty, the 

repeatability suggests that the comparison is accurate within 

0.6 percent. The loss in adiabatic efficiency shows a 

significant improvement from previous over-the-rotor acoustic 

treatment designs, therefore, the acoustic benefit of the 

treatment will be evaluated in the 9x15 Low-Speed Wind 

Tunnel at NASA Glenn Research Center. Thermocouples in 

the acoustic treatment showed a rapid increase in treatment 

temperature when the blade tip speed approached the speed of 

sound. In order to evaluate the effectiveness of an over-therotor acoustic treatment on a fan with supersonic tip speeds, 

the treatment’s temperature limit should be above 250 °F 

(121 °C).

References

1. Sutliff, Daniel L., and Jones, Michael G., “Foam-Metal 

Liner Attenuation of Low-Speed Fan Noise,” AIAA 

2008-2897, AIAA Aeroacoustics Conference, Vancouver, 

BC, May 5-7, 2008.

2. Elliott, David M., Woodward, Richard P., Podboy, Gary 

G., “Acoustic Performance of Novel Fan Noise Reduction 

Technologies for a High Bypass Model Turbofan at 

Simulated Flight Conditions,” AIAA 2009-3140, AIAA 

Aeroacoustics Conference, Miami, FL, May 11-13, 2009.

3. Hughes, Christopher E., Gazzaniga, John A., “Effect of 

Two Advanced Noise Reduction Technologies on the 

Aerodynamic Performance of an Ultra High Bypass Ratio 

Fan,” AIAA 2009-3139, AIAA Aeroacoustics 

Conference, Miami, FL, May 11-13, 2009.

4. Sutliff, Daniel L., Elliott, Dave M., Jones, Michael G., 

and Hartley, Thomas C., “Attenuation of FJ44 Turbofan 

Engine Noise with a Foam-Metal Liner Installed Overthe-Rotor,” AIAA 2009-3141, AIAA Aeroacoustics 

Conference, Miami, FL, May 11-13, 2009.

5. Bailey, Everett E., “Effect of Grooved Casing Treatment 

on the Flow Range Capability of a Single-Stage Axial 

Flow Compressor,” NASA TM X-2459, Jan., 1972.
NASA/TM—2013-218066 8

6. Hathaway, Michael D., “Passive Endwall Treatments for 

Enhancing Stability,” NASA/TM—2007-214409, Jul., 

2007.

7. Van Zante, Dale E., Podboy, Gary G., Miller, Christopher 

J., Thorp, Scott A., “Testing and Performance 

Verification of a High Bypass Ratio Turbofan Rotor in an 

Internal Flow Component Test Facility,” GT2007-27246

or NASA/TM—2009-215661, Turbo Expo 2007, 

Montreal, QC, May 14-17, 2007.

8. Hughes, Christopher E., Jeracki, Robert J., and Miller, 

Christopher J., “Fan Noise Source Diagnostic Test –

Rotor Alone Aerodynamic Performance Results,” AIAA 

2002-2426 or NASA/TM—2005-211681, AIAA 

Aeroacoustics Conference, Jun. 17-19, 2002.

9. Nark, Douglas M., and Jones, Michael G., “Broadband 

Liner Optimization for the Source Diagnostic Test Fan,” 

AIAA 2012-2195, AIAA Aeroacoustics Conference, 

Colorado Springs, CO, Jun., 4-6, 2012.

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The Aerodynamic Performance of an Over-the-Rotor Liner With Circumferential Grooves on 

a High Bypass Ratio Turbofan Rotor 

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13. SUPPLEMENTARY NOTES

14. ABSTRACT

While liners have been utilized throughout turbofan ducts to attenuate fan noise, additional attenuation is obtainable by placing an acoustic 

liner over-the-rotor. Previous experiments have shown significant fan performance losses when acoustic liners are installed over-the-rotor. 

The fan blades induce an oscillating flow in the acoustic liners which results in a performance loss near the blade tip. An over-the-rotor liner 

was designed with circumferential grooves between the fan blade tips and the acoustic liner to reduce the oscillating flow in the acoustic 

liner. An experiment was conducted in the W-8 Single-Stage Axial Compressor Facility at NASA Glenn Research Center on a 1.5 pressure 

ratio fan to evaluate the impact of this over-the-rotor treatment design on fan aerodynamic performance. The addition of a circumferentially 

grooved over-the-rotor design between the fan blades and the acoustic liner reduced the performance loss, in terms of fan adiabatic 

efficiency, to less than 1 percent which is within the repeatability of this experiment. 

15. SUBJECT TERMS

Over-the-rotor; Turbofan; Performance 

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