Effect of Tip Clearance on Fan Noise and Aerodynamic Performance

Transcription

1 11th AIAA/CEAS Aeroacoustics Conference (26th AIAA Aeroacoustics Conference) May 2005, Monterey, California AIAA Effect of Tip Clearance on Fan Noise and Aerodynamic Performance Christopher E. Hughes * Richard P. Woodward Gary G. Podboy NASA Glenn Research Center, Cleveland, OH The design of effective new technologies to reduce aircraft propulsion noise is dependent on identifying and understanding the noise sources and noise generation mechanisms in the modern turbofan engine, as well as determining their contribution to the overall aircraft noise signature. Therefore, a comprehensive aeroacoustic wind tunnel test program was conducted as part of the NASA Quiet Aircraft Technology program called the Fan Broadband Source Diagnostic Test. The test was performed in the anechoic NASA Glenn 9- by 15-Foot Low Speed Wind Tunnel using a 1/5 scale model turbofan simulator that represents a current generation high bypass turbofan engine. The investigation was focused on the simulated bypass section of the turbofan engine. The technical objectives of the test were to: identify the noise sources within the model and determine their noise level; investigate several component design technologies by evaluating their impact on the aerodynamic and acoustic performance; and conduct detailed flow diagnostics within the research model to help understand the physics of the flowfield. This report will present the results obtained for one aspect of the test - the effect of the fan tip clearance on the bypass fan and stage performance. The aerodynamic performance, farfield acoustics, and Laser Doppler Velocimeter measurements obtained for the fan with four different fan tip clearances are shown. The fan tip clearances investigated were designated at the closest approach point of the fan tip to the rubstrip casing over the fan at 100% fan speed. These fan tip clearances were; nominal, or line-to-line (0.000 inch), representing a new engine;.020 inch, represent a typical older engine about to be removed from service for maintenance; inch, representing the clearance after the aircraft experiences hard maneuvering or a hard landing; and inch, representing the clearance after a severe aircraft landing. The aerodynamic performance results indicate that the fan adiabatic efficiency was highest with nominal tip clearance (about 91.8% at design speed) and decreased as fan tip clearance increased. Differences in fan efficiency between tip clearance configurations was small below 77.5% fan design speed, about 0.5% maximum, and got larger as fan speed increased, to around 1% maximum at 100% fan design speed. The decreases in efficiency are due to a lower blade loading and higher temperature rise over the outer 20% of the blade span; the mechanism is the tip leakage flow from pressure to suction surface of the fan blade, which increases as the fan tip gap increases. Farfield acoustic results show that changes in the noise level are primarily aft-radiating and are on the order of 1 to 5 db in fan broadband noise level as measured for the rotor-alone configuration (which had no stators). There were very small changes in noise level with tip clearance for the 54-vane Baseline Outlet Guide Vane configuration. Small changes in noise were also seen for the other two Outlet Guide Vane configurations tested - the 26-vane radial Low Count Outlet Guide Vanes and the 26-vane aft swept Low Noise Outlet Guide Vanes. The rotor alone acoustic results suggest that tip clearance changes induce broadband noise changes at the fan tip, but that the noise differences are masked by the rotor wake/stator interaction noise generated with the Outlet Guide Vanes installed. The magnitude of the rotor/stator interaction tones showed small increases with increasing fan tip gap at fan speeds below transonic, indicating slightly stronger interactions with the larger tip vortex wakes that formed with larger tip gaps. The broadband levels, for the most part, only showed the effect of Outlet Guide Vane geometry on the magnitude of the noise as the tip gap changed. The Laser Doppler Velocimetry flow field results show that the flow downstream of the tip of the blades changes very little with changes in the tip clearance when operating at the approach condition. At both the cut-back and take-off speeds, significant changes in the tip flow occur with changes in the tip clearance. Since these changes in the tip flow are not * Senior Research Engineer, RTA/Acoustics Branch, Brookpark Rd/MS Senior Research Engineer, RTA/Acoustics Branch, Brookpark Rd/MS 54-3, AIAA Member. Senior Research Engineer, RTA/Acoustics Branch, Brookpark Rd/MS 54-3, AIAA Member. 1 This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States.

2 accompanied by significant changes in the noise level, it appears that this noise source the tip clearance flow is masked by the OGVs, as evidenced by the far field acoustics by comparing rotor alone results and results obtained with Outlet Guide Vanes installed. Nomenclature A Cross sectional area, in 2 c Airfoil chord, in M Mach number N Mechanical fan speed, RPM N c Corrected fan speed, N θ NP c Percent of corrected fan design speed, Nc 100, % N dp P Pressure, psia R Gas constant, ft lbf/lbm R T Temperature, R t Airfoil thickness, in γ Specific heat ratio, 1.4 δ P t,o Pressure correction to standard day conditions, γ 1 η P t γ 1 P t,o Adiabatic efficiency, T t 1 T t,o θ ω ω c Θ Temperature correction to standard day conditions, Weight flow rate, A bm γ R Corrected weight flow rate, ω Acoustic angle, deg T t,o Pt,bm M bm γ + 1 Tt,bm γ M 2 bm θ δ, lbm/sec ( γ 1), lbm/sec Subscripts ae Adiabatic efficiency bm Bellmouth inlet condition c Corrected condition dp Design point em Emission f Fan value f Force geom Geometric m Mass o Freestream condition s Static condition st Stage value t Total condition 2

3 I. Introduction N recent years, commercial aircraft noise has become a major concern for aircraft owners and airport operators. I The increased frequency of takeoffs and landings has produced an increasing number of complaints from local residents. The Federal Aviation Administration in the United States and the International Civil Aviation Organization, the international organization that regulates environmental noise issues, have responded to these complaints by issuing increasingly more stringent noise regulations and curtailed flight operations for aircraft, forcing aircraft and engine manufacturers to pursue quieter aircraft designs. With the support of Congress, NASA and the U.S. aircraft and engine manufacturing companies have joined to cooperatively investigate high-risk technologies for reducing aircraft noise through programs such as the NASA Advanced Subsonic Technology (Ref. 1-4) and Quiet Aircraft Technology Programs. As part of an overall NASA effort to reduce aircraft noise, technical programs were initiated starting in 1989 with the major U.S. aircraft manufacturers to investigate noise reduction technologies. NASA established aggressive goals aimed at reducing the noise signature of 1992 technology turbofan engines by 6 EPNdB (Effective Perceived Noise db) by the year Engine studies were conducted across a wide range of engine and aircraft operating cycles to identify and quantify the benefit of potential noise reduction concepts. Several noise reduction technology concepts were investigated using scale model wind tunnel testing of turbofan engine simulators, and the noise reduction potential successfully demonstrated in most cases (Ref. 5-10). However, new noise reduction standards which were driving new noise reduction program goals were aggressively pushing the technology. The goals of the NASA Quiet Aircraft Technology Program, the follow-on to the Advanced Subsonic Technology program, sought to reduce turbofan noise by another 4 EPNdB by 2006, using 1997 High Bypass turbofan engine technology as the baseline. To achieve this new noise reduction goal, novel technical approaches to noise reduction technology for turbofan engines would be necessary. Therefore, in order to more fully understand the noise sources and noise generation mechanisms in a modern turbofan engine and be able to properly guide further noise reduction technology development, a comprehensive scale model wind tunnel test of a turbofan simulator was planned, called the Fan Broadband Source Diagnostic Test, or just Source Diagnostic Test (SDT). The test was a cooperative effort between NASA and General Electric Aircraft Engines. It was a two-phase experimental investigation conducted in the NASA Glenn anechoic 9- by 15- Foot Low Speed Wind Tunnel (9x15) to identify and understand the noise source mechanisms within a turbofan engine and determine their individual contributions to the overall engine system noise signature. For this test, the bypass stage portion of a medium pressure ratio, high bypass ratio turbofan engine was simulated in approximately 1/5 model scale. The fan model consisted of a 22-inch diameter, 22-blade, wide-chord fan, an outlet guide vane or stator assembly, and a simulated flight-type nacelle including a fixed area nozzle and an inner flowpath contour which simulated the outer core cowling. The emphasis for this test was placed only on simulating the bypass stage portion of the engine, not the booster core or power stage, eliminating the possibility of contaminating the fan noise field with the noise from a simulated core. Likewise, to minimize the noise contamination sources within the fan model, the outlet guide vanes were used to support the nacelle, eliminating any struts or pylons from the flowpath. The fan model was powered by the NASA Glenn Ultra High Bypass (UHB) Drive Rig propulsion simulator. In order to simulate aircraft flight effects during takeoff roll, approach and landing phases, wind tunnel velocities up to Mach 0.10 were provided during acoustic testing. The Source Diagnostic Test had several technical objectives related to farfield acoustics, aerodynamic performance, and fan flow field diagnostics, and was completed in two phases. The first phase of testing, completed in 2000, verified the usability of the proposed fan as a baseline for future technology development. The aerodynamic performance and farfield acoustics of the fan were measured. In addition, the effect of the number of outlet guide vanes (OGVs) used, as well as the effect of aft radial sweep and chord length, on the aerodynamic performance and noise level of the fan model was investigated with three different outlet guide vane designs (Ref. 11). The overall test had several areas of investigation including acoustic mode measurements using sensors located on the inner surface of the fan duct, spinning mode measurements using a rotating acoustic rake in both the inlet and the nozzle, unsteady surface pressure measurements on two different types of outlet guide vanes, and detailed flow diagnostic measurements using Laser Doppler Velocimetry (LDV) and hot-wire anemometry (Refs ). Also completed during this test phase was the first rotor alone aero-acoustic testing in a wind tunnel using realistic 3

4 simulation of the bypass flow line geometry (Ref. 14), including the fan tip clearance. This part of the testing was important for understanding the contribution of the fan noise sources to the overall engine system noise. In phase two, completed in 2003, the same fan model hardware components were used to determine the effect of fan tip clearance and bypass nozzle exit area on the aerodynamic performance and noise. The results presented in this paper have a direct relationship to current turbofan engine technology, and can be scaled to full-scale aircraft engine applications. In addition, a comprehensive set of flow diagnostic surveys using LDV were performed to measure the unsteady velocity and turbulence components within the internal model flowfield. These data allow a better understanding of the flow physics and how the flow interacts with the model hardware, leading to identifying and understanding the mechanisms that produce noise within the model. This paper will discuss the test results to determine the effect of fan tip clearance on the fan and bypass stage aerodynamic and acoustic performance. Four different fan casing rubstrip geometries were designed and fabricated for use with the baseline fan and outlet guide vanes. Fan tip clearances of nominal (or line to line), inch (.020 ), inch (.030 ) and inch (.040 ) were investigated. The fan tip clearance designations represent the mean clearance between the fan tip and the rubstrip casing at the fan design speed (100% corrected fan speed). The nominal clearance was defined by GE as representative of their design clearance for newly manufactured engines. NASA Glenn used the.020 clearance as the baseline for all acoustic data, since that clearance simulated the noisiest configuration that would be used in operation. The aerodynamic performance of the fan and OGVs with each of the four fan casing configurations will be discussed in terms of fan and stage pressure, temperature and adiabatic efficiency maps including radial profiles downstream of the fan. Farfield acoustic data for the fan model with three fan casing configurations (nominal,.020 and.030 ) will be presented in terms of the Overall Sound Power Level and Effective Perceived Noise Level at various fan speeds. The acoustic results will also address the effect of OGV design (Baseline, Low Count and Low Noise) on the fan tip clearance results. Finally, flow diagnostics will be shown which will present the mean velocity and turbulence components in the fan wake for the four fan tip clearances at several fan speeds, showing the changes in flow and turbulence with increasing fan tip clearance. II. Research Apparatus A. Facilities The SDT was conduced in the NASA Glenn 9- by 15-Foot Low Speed Wind Tunnel (9x15). The anechoic facility is operated as an open loop, continuous flow wind tunnel at atmospheric pressure conditions and is capable of producing Mach numbers from up to 0.23 (Ref. 17), and can be used for static propulsion system testing as well. Flow conditioning upstream of the test section allows the facility to produce very low freestream turbulence and distortion levels, making it ideal for acoustic testing of propulsion systems (Ref. 18). The test section acoustic treatment is capable of absorbing sound reflections down to 250 Hz (Ref ). The NASA Glenn Ultra High Bypass (UHB) Drive Rig was used to power the model fan test article. This propulsion simulator is very similar to one used by General Electric Aircraft Engines (GEAE) Universal Propulsion Simulator (Ref. 22). A four-stage air turbine generates the power that is supplied to the fan model through a common shaft connection. The air turbine is driven by high pressure (up to 350 psia), high temperature (up to 550 F) air that is supplied to it from tubes running through a support strut that mounts the UHB Drive Rig in the wind tunnel test section. The UHB Drive Rig can generate a maximum of 5,000 shaft horsepower at 16,850 RPM. Figure 1 is an overhead view of the 9x15 test section showing the fan model hardware and UHB Drive Rig in relationship to the farfield acoustic measurement system and to the wind tunnel test section. Figure 2 is a cutaway view of the UHB Drive Rig and a table of maximum performance parameters. B. Fan Module The research model, or fan module, used was a 1/5-scale model representation of the bypass stage of a current generation high bypass turbofan aircraft engine. The fan module was designed and built by GEAE with partial funding under contract to NASA Glenn. Only the bypass section of the engine was simulated in order to ensure that the fan model noise field was not contaminated by noise from a core section simulator. The fan module consisted of the fan, the outlet guide vanes (OGVs) and a flight-type nacelle. The nacelle included a flight-type inlet, a cowl and a fixed-area, flight-type bypass exhaust nozzle. In order to minimize adding additional noise sources within the model, the OGVs were designed to provide structural support for the nacelle, thereby eliminating the need for struts 4

5 normally present in a turbofan engine. In addition, neither pylon nor bifurcations were simulated in the model. Figure 3 shows a cutaway view of the fan module in the flight configuration, showing the location of various model components. The fan used for this test was 22 inches in diameter and had 22 individual, wide-chord, titanium blades. In combination with the 54 vane baseline OGVs, the fan model had a stage design point pressure ratio of 1.47 at a model corrected speed of 12,657 N c, corresponding to a design point fan tip speed of 1,215 feet per second. Table 1 provides a summary of the design parameters for the fan. The fan was a scale model designed and previously tested by GEAE, who designated the fan as R4. The fan was originally designed to operate in conjunction with a powered core simulator. As a result, the performance level at its design point could not be achieved in this test since this installation did not include a core simulation. However, since this fan was meant to represent current turbofan engine technology, the performance compromise was deemed acceptable for this test. The fan was tested with four different rubstrip casings that were designed to vary the fan blade tip clearance at the fan design point (100% corrected fan speed, or 12,657 N c ). The four tip clearances were: nominal, representing zero clearance between the fan blade tip and the flowpath surface at the fan design speed, and the GEAE design tip clearance for a new engine;.020, representing the clearance of an engine near a maintenance cycle;.030, representing an engine after a very hard landing; and a.040, representing an engine after hard maneuvering or a severe landing. All four fan rubstrip casings had smooth contours on the flowpath surface that were blended into the flow lines ahead of and behind the rubstrip area. The.020 clearance was selected as the baseline configuration, since it could potentially simulate the noisiest engine configuration under normal service conditions. As part of the fan module design, there were three distinct OGV designs, representing an acoustic baseline and two additional configurations designed to reduce the level of specific noise sources in the fan module. These three OGV designs were: Baseline with 54 narrow chord, high aspect ratio, radial vanes; Low Count with 26 wide chord, low aspect ratio, radial vanes; and Low Noise with 26 wide chord, low aspect ratio, aft swept vanes. A complete description with diagrams and photographs of the three OGV designs and acoustic design considerations, along with the aerodynamic and acoustic performance measured with the R4 fan are given in Reference 11. In order to maintain the aerodynamic loading for each OGV configuration, the solidity between the three designs was held nearly constant. In addition, the flowpath geometry was designed to achieve as close to the same flow velocity as possible between the three designs. Figure 4 is a schematic cutaway illustrating the relative orientation of the three OGV configurations in the fan module. Table 2 is a summary of the design parameters for the three OGV configurations. To establish the fan and OGV performance, the fan module installation included a uniform-inflow bellmouth inlet and either a fixed-area flight-type nozzle or a Variable area Fan Exit Nozzle (VFEN). The fixed area nozzle was used to obtain the fan sea level performance on a representative turbofan engine operating line. The VFEN was used to obtain fan and stage performance over a range of fan speed operating conditions and simulated aircraft flight conditions. The fan operating point was changed by varying the exit area, and hence the back pressure on the fan and weight flow through the fan stage, while at a constant fan speed. A change in the fan back pressure simulates a change in the aircraft flight speed and altitude. In Figure 5a, a photograph of the fan module in the aerodynamic performance configuration in the 9x15 test section are shown. Figure 5b is a close-up view of the VFEN arrangement. III. Research Instrumentation A. Aerodynamic Wind tunnel operating conditions, which were used as freestream conditions during acoustic testing, were determined using a ceiling mounted pitot-static/thermocouple rakes. Fan model reference conditions were determined using a floor mounted, cruciform-shaped rake located near the fan centerline and upstream of the bellmouth inlet. The fan weight flow was obtained with a calibrated bellmouth inlet which used static pressure measurements and a flow correlation function relating the average bellmouth static pressure to the fan weight flow. Fan and stage performance were determined using fixed total pressure/total temperature rakes and surface static pressures on the inner and outer flow path, located behind the fan and OGVs. Fan performance was measured using three rakes and stage performance was measured with seven rakes. Each rake consisted of seven measurement 5

6 sensors, and each sensor contained a total pressure probe and a total temperature probe co-located within an aspirated stagnation tube. The sensors on each rake were located radially in such a way as to provide flow conditions at the center of equal areas. In addition, surface mounted static pressures were located at several axial locations in the fan module for calculating internal velocities. During the fixed operating line testing with the fixed area nozzle installed, only fan performance could be measured, since the OGV performance rakes not designed to be installed with the fixed nozzle configuration. A complete description of the aerodynamic performance instrumentation used in this test can be found in Reference 11. Forces produced by the fan and OGV were measured using two different types of force balances. Thrust and torque produced by the fan was measured with a two-component rotating force balance. Signals from this balance were transmitted to the data system using a 104 channel slip ring system internal to the UHB Drive Rig. Thrust and drag forces produced by the OGVs (including the nacelle and inner flowpath hardware downstream of the fan) were measured using a six-component static force balance mounted directly on the UHB Drive Rig. Signals were sent directly to the data system through wiring channels located in a fairing mounted in front of the Drive Rig strut. Figure 3 shows the location of the balances and their physical relationship to the fan module components. The accuracy of the measured thrust is ±10 lbf., or ±0.25% of the full scale measurement range (4000 lbf.) of the combined balances. A complete description of the force balance components and technique used in measuring performance of turbofan simulators in wind tunnel testing at Glenn using can be found in Reference 23. B. Acoustic Figure 6a is a photograph of the fan module in the acoustic, flight-type configuration installed in the 9x15. In Figure 6b, the downstream fixed microphones and sideline traversing microphone probe can be seen next to the fan model on the left. Sideline acoustic data were acquired with a computer-controlled translating microphone probe and with three aft microphone assemblies mounted to the tunnel floor. The translating microphone probe acquired data at 48 sideline geometric angles from 27.2 o to o relative to the fan rotor plane. The translating probe traverse was 89 inches from the fan rotational axis (about four fan diameters). A fixed, wall-mounted microphone probe was placed adjacent to the translating probe in the home position (134.6 o, maximum aft travel) as a reference for the translating microphone. The three fixed microphone assemblies were mounted at the home axial position to acquire aft acoustic data at geometric angles of 140, 150, and 160 o. The acoustic data were acquired through a digital computer system and stored for post-run analysis. C. LDV To conduct LDV wake surveys within the fan module, it was necessary to place part of the LDV system inside the test section of the wind tunnel. Figure 7a shows a photograph of the LDV traverse system located on the side of the fan model. The traverse was used to move the LDV probe volume radially and axially relative to the model. The LDV system optics are located behind the cylindrical shield shown in the photo. This shield was installed to keep the tunnel flow from striking the optics. Figure 7b shows a photograph taken with the cylindrical shield removed. In this photo the fiber optic cables used to deliver the laser beams into the tunnel, the transmitting optics used to direct the beams into the model, and one set of receiving optics can be seen. The LDV system is a four-beam, two-color, backscatter system which allows the measurement of two components of velocity simultaneously. Two green beams were used to measure the axial component of velocity, while two blue beams allowed the measurement of the tangential component. The photo provided in Figure 7b shows one of two optical arrangements used during the test. As pictured, only one component of velocity could be measured at a time. Initially, another optical arrangement, one employing two sets of receiving optics (one above and one below the transmitting optics) was used to conduct the wake surveys. During the initial surveys it was possible to measure both the axial and tangential velocity components simultaneously. The LDV flow diagnostic wake surveys were conducted with bellmouth inlet installed on the model. The bellmouth inlet provided essentially the same flow into the fan as the flight inlet, but allowed the tunnel to be run at a lower speed. This provided the following benefits: 1) the test was conducted at a lower cost and 2) the LDV system components mounted in the test section were subjected to lower air loads. Two windows installed in the side of the model permitted optical access to the internal flow. These two windows are shown in the photograph of Figure 7b. The downstream window shown at the left was used to acquire the wake surveys. These windows, made 6

7 of 0.1 in. thick sodium alumino silicate, were slumped in a furnace to have the same flow line contour shape as the model at the fan case. In order to acquire LDV data at the axial location corresponding to the leading edge of the Baseline OGVs, the Low Noise OGVs were installed in the model. Figure 8 shows the location at which LDV wake measurements were made relative to the fan module hardware. The survey plane was located 3.12 fan chords downstream of the nonrotating position of the fan tip trailing edge. The increase in the axial length between the fan trailing edge and the OGV leading edge provided by the swept OGVs allowed a large downstream LDV window to be used, which in turn allowed the Baseline OGV leading edge axial location to be viewed by the LDV system. Test results previously obtained for the fan module during the SDT Phase 1 have shown that the fan and stage performance were relatively unaffected by replacing the Baseline OGVs with the swept OGVs (Ref. 11). IV. Experimental Procedure A. Aerodynamic Performance Fan and stage aerodynamic performance maps were obtained with the baseline R4 fan and three of the fan tip clearance rubstrip casings nominal,.020 and.030 and with the Baseline OGVs installed. In addition, the fan and stage performance on the fixed area nozzle operating line was obtained with the R4 fan, Baseline OGVs and all four fan tip clearance configurations. A Mach number of.05 was set in the test section during aerodynamic testing in order to provide uniform temperature and pressure distributions into the fan model as well as prevent the fan from creating and ingesting vortices from the test section surfaces. To eliminate the day-to-day variations in pressure and temperature that affect the performance calculations, the fan and stage performance parameters were corrected to standard day pressure and temperature conditions, where required. To insure that data was acquired at steady state conditions, a 30 second settling time interval was maintained after each new fan operating condition was reached. In addition, pressure and temperature information from the data system was time averaged over a 10 second sampling. 1. Fan and Stage Mapping Fan and stage (or OGV) performance mapping was conducted with the bellmouth inlet and the VFEN installed on the fan module. A fan speed range from 50% to 100% of the fan design speed was investigated. At each speed condition, the maximum weight flow was achieved at the maximum nozzle area with the VFEN fully open. To determine stage adiabatic efficiency, it was assumed that there was no loss in total temperature across the OGVs; this allowed the total temperature data from the fan rakes to be used in the stage performance calculations. This method is a more accurate determination of stage adiabatic efficiency since variations in temperature measurements between the fan rakes and the stage rakes caused by instrumentation variations are eliminated; as a result, any accompanying errors in the stage adiabatic efficiency calculation caused by the instrumentation variations are eliminated. Overall values for the fan and stage performance were obtained by averaging the seven radial profile values for each performance parameter. To minimize the risk of damaging the fan blades by an unintentional hard fan rub, the fan stall region located at the lower fan weight flow conditions was intentionally avoided. With this fan design, an approaching stall condition was indicated by an increase in the fan blade stress. Therefore, the minimum fan weight flow was established when the fan blade stress measured with blade mounted strain gauges reached a predetermined limit, typically 25% of the fan material yield strength. This limit, however, varied with fan speed. The 87.5% speed line was unusually sensitive (higher blade stress at higher weight flow conditions), so the minimum weight flow boundary at this speed was larger than the minimum weight flow at the other fan speed lines. The shape of the maps had a distinct bend, or knee, because of this sensitivity. 2. Fixed Nozzle Operating Line Fan performance on the fixed area nozzle operating line at sea level conditions was obtained using the bellmouth inlet and the fixed area, flight-type nozzle. However, only the fan performance could be obtained in this hardware configuration since the stage performance rakes could only be installed with the VFEN installed. Therefore, once the fan weight flow and fan operating parameters were established for the fixed nozzle operating line, the corresponding stage performance was obtained with the VFEN installed by adjusting the nozzle exit area to match the fixed nozzle fan pressure and weight flow at each corresponding fan speed. 7

8 B. Farfield Acoustics The farfield acoustic testing was conducted with only three fan tip clearance configurations nominal,.020 and.030. The.040 tip clearance was not tested acoustically because this clearance did not represent an engine configuration that would be routinely flown on an aircraft, and also due to time constraints in the test schedule. In addition, the tip clearance variations were tested with all three OGV configurations and rotor alone using the NASA Glenn Rotor Alone Nacelle (RAN) (Ref. 14). It has been shown that the rotor alone aerodynamic performance is very close to the fan and OGV aerodynamic performance (Ref. 14). All of the fan and stage acoustic data were acquired at a tunnel Mach number of 0.10, which was sufficient to achieve acoustic flight effect (Ref. 24). Sideline data are presented in terms of emission angles. The emission angles are related to the geometric or observed angles by the relationship: Θ em = Θ geom sin -1 (M o sin Θ geom ) where Θ em and Θ geom are, respectively, the emission and observed sideline angles, and M o is the test section Mach number. At Mach 0.10 then, the observed angles for the sideline translating microphone probe range from 25 o to 130 o, and the three fixed microphones measure aft observed angles of 136, 147, and 158 o. This angular range was sufficient to define the sideline noise profile for this aft-dominated fan stage for subsequent Effective Perceived Noise Level (EPNL) calculations. Digital acoustic data were processed as constant bandwidth spectra. Spectra were acquired and averaged at each translating probe or fixed microphone position with 5.9 and 59 Hz bandwidths. These constant bandwidth spectra were electronically merged and used to generate 1/3-octave spectra, using the 5.9 Hz bandwidth results for lower 1/3 rd octave frequencies, and the 59 Hz bandwidth results for the higher frequencies. Sound Power Level (PWL) spectra were calculated from the Sound Pressure Level (SPL) spectra assuming spherical symmetry through the range of sideline data acquisition. Fan stage EPNL was calculated for a 1500 ft. flyover (at 0.10M) and a 3.35 scale factor. Although this scale factor is different than the scale factor for the model hardware, that difference does not affect the noise results. Possible noise contributions outside the sideline range were ignored. C. LDV surveys Wake data were acquired at four corrected rotor speeds, 6,329, 7,808, 11,074, and 12,657 (corresponding to 50%, 61.7%, 87.5% and 100% of the corrected fan design speed). The 7,808, 11,074, and 12,657 speeds correspond to the acoustic rating points of approach, cutback, and sideline or takeoff conditions, respectively, for the R4 fan. The wake surveys were conducted to determine how the wake flows vary with fan tip clearance and rotor speed. The tunnel flow was seeded with polystyrene latex (PSL) spheres that were manufactured at the NASA Glenn Research Center. The nominal size of the PSL spheres is estimated to be approximately 0.7 micron in diameter. The polystyrene spheres are diluted with ethanol and sprayed into the wind tunnel using a set of nine spray nozzles located approximately 80 feet upstream of the test section. The liquid solvent evaporates by the time it reaches the test section, leaving behind the solid spheres on which the LDV data is obtained. The test section Mach number during the LDV testing was approximately 0.05, just fast enough to move the seed particles downstream in the wind tunnel to the fan model. The individual velocity measurements were sorted into circumferential bins around the rotor using shaft angle encoders fed with the once-per-revolution signal of the rotor. These encoders segmented the 360 degrees of rotor revolution occurring between two consecutive once-per-revolution pulses into 1100 bins of equal width (50 bins per blade passage). Each time a velocity measurement was made, the encoder output was sampled to determine the number of bins generated since the occurrence of the previous once-per-rev pulse. The velocity and corresponding bin number were then stored in the computer as a data pair. Data were acquired at the survey measurement location over many rotor revolutions until either a preset number of measurements had been acquired on one of the two LDV channels, or until the maximum time allotted for the data acquisition had elapsed. On-line data plots were used to determine the number of measurements required to accurately resolve the flows occurring within the individual blade passages. In general, the higher the unsteadiness in the flow, the greater the number of measurements required to resolve the flow. On average, more than 40,000 velocity measurements per component were obtained at each combination of measurement location and operating condition. A more complete description of the LDV system and technique as used for flow diagnostics in turbofan simulators during wind tunnel tests at Glenn can be found in Reference 12. 8

9 V. Experimental Procedure A. Aerodynamic Performance For the results presented in this section, the precision of the performance calculations is based on empirical observation and repeat data points. The accuracy of the data acquisition systems used during testing were ± psia for pressure and ± 0.25 F for temperature. However, the data systems were configured to provide timeaveraged measurements at a high sample rate. For temperature and pressure, the data values are based on an average of ten, one-second averages, with each one-second average based on the average of 20,000 samples. Therefore, the precision of the discrete performance points is higher than the results based on discrete data samples: for pressure ratio, the precision is ±0.0003; for temperature ratio, the precision of the results is ±0.001; and for adiabatic efficiency, the precision of the results is ±0.25%. Fan performance maps are presented in Figure 9. The corrected weight flow, fan total pressure ratio, total temperature ratio and adiabatic efficiency are shown in Figures 9a through 9d, respectively, for three fan tip clearances - nominal (.000 ),.020,.030. Also shown on each of the performance map plots is the operating line performance as measured for all four fan tip clearances nominal (.000 ),.020,.030 and.040. As stated in an earlier section, the last data point on the far left of each speed line on the plot does not represent the fan stall line. Instead, this minimum weight flow condition data was obtained at what was considered a safe operating distance away from the fan stall line in order to avoid a fan tip rub or any possible damage to the fan blades due to high blade stress caused by unintentionally entering a fan stall condition. The results in Figure 9 show that increasing the fan tip clearance produces minor changes in weight flow, pressure ratio and temperature ratio below 87.5% corrected fan speed. Above that speed, minor reductions in pressure ratio and somewhat larger increases in temperature ratio can be seen. The most obvious differences in performance can be seen in the fan adiabatic efficiency map in Figure 9d. At the lower fan speeds and lower weight flow conditions, differences of about 0.5% can be seen from nominal to.030 clearance. Moving toward higher weight flow on each speed line, the performance differences become larger. On the 60% speed line, the.030 tip clearance shows a significant loss in performance (from 0.75% to 1% from nominal to.030 clearance) as the weight flow increases. The.030 tip gap is obviously large enough to allow significant flow leakage at the fan tip which that is causing the drop in efficiency as the weight flow and axial velocity component increase. The.030 clearance shows larger losses than the other two clearances on the other higher fan speed lines at higher weight flow conditions, but the loss at 60% seems more dramatic. Some of this loss may also be due to larger error bands on the lower speed lines with the lower pressure ratio, lower temperature ratio. At higher fan speeds, the.020 clearance also begins to show a drop in efficiency as the weight flow increases. The largest differences in efficiency between the nominal,.020 and.030 tip clearances can be seen near the fan design point at 100% corrected speed. Here, the loss in peak efficiency is about 0.5% from nominal to.020 clearance (from 93.6% to 93.1%), and 1.1% from nominal to 0.30 clearance (from 93.6% to 92.5%). In the same figure (Fig. 9d), the performance differences between the four fan tip clearance configurations along the fixed area nozzle operating line are shown. At fan speeds up to 77.5%, the range of change in adiabatic efficiency is about 0.5% from nominal to.040. Above this speed, the lines start to diverge as the larger tip clearances allow some loading changes at the fan tip. Interestingly, the lines seem to merge into two distinct groups. The fan efficiency differences for the nominal and.020 clearances are small, less than 0.25%. Likewise, the.030 and.040 differences are also small, again less than 0.25%. However, the two groups are separated by about 0.50% in efficiency. At the maximum fan speed of 100% corrected, where the largest differences in performance occur, the.030 and.040 clearance efficiency differences are still small (less than 0.25%). The.020 clearance performance also falls off compared to the nominal. At this speed, the fan adiabatic efficiency for the nominal clearance is 91.9%, the.020 is 91.5%, the.030 is 91.1% and the.040 is 90.9%. The difference in corrected weight flow for this speed ranges from 97.2 lbs/sec to 97.4 lbs/sec from.040 to nominal clearance, respectively. It appears that differences in fan performance on the fixed area nozzle operating line only become significant as the tip clearance increases beyond.020 at all fan speeds, but especially above 77.5%. At fan speeds above 95% corrected, more significant differences occur at any clearance larger than the nominal. Beyond the.030 fan tip clearance, it appears that no further significant changes in performance occur. This indicates that no further drop in blade loading at the fan tip occurs, and tip flow leakage has stabilized. 9

10 In Figure 10, radial profiles of the fan pressure ratio, temperature ratio and adiabatic efficiency measured with the fixed fan exit rakes are shown. In Figure 10a, the change in total pressure ratio and blade loading at the fan tip with changes in tip clearance are presented. Loss in performance is significant only at the furthest outboard rake sensor location near the fan tip and only at 100% fan speed. The total temperature profiles in Figure 10b also show the drop in performance as indicated by the increase in total temperature ratio with increasing tip clearance, but again it appears only at the furthest outboard sensor near the fan tip. These differences become more pronounced starting at 87.5% fan speed, which is a lower fan speed than was shown for significant differences in the total pressure ratio profiles (Fig. 10a). This result seems to indicate that total temperature is more sensitive than total pressure to the fan tip clearance. The lack of performance change as the tip clearance increases from.030 to.040 in both Figures 10a and 10b can be seen, and especially for temperature ratio, where there seems to be no difference in performance between these tip clearances. The fan adiabatic efficiency radial profiles are shown in Figure 10c. With a couple of exceptions, the differences in efficiency with changes in tip clearance only appear in the data at the furthest outboard sensor near the fan tip. Differences in efficiency at this location range from about 4.2% at 100% fan speed to about 1.2% at 61.7% fan speed as the clearance increases from nominal to.040. These results verify the previous performance profile results for pressure and temperature ratio in Figures 10a and 10b, respectively, which is that the largest performance differences are between nominal and.020, and.020 to.030 clearances. The difference between the.030 and.040 clearances is almost constant at all speeds, about 0.2%, indicating that the efficiency penalty levels off. The pronounced jump or increase in efficiency of about 2% in the profiles at 87.5% fan speed, between 60% and 72% blade span, are a result of flow interaction with the shocks at that blade span location. The fan is transitioning to sonic conditions at 87.5% fan speed, and as a result a change in temperature rise occurs across the shock. The stage performance maps showing the differences in total pressure ratio and adiabatic efficiency with changes in tip clearance with the Baseline OGVs are presented in Figures 11a and 11b, respectively. As discussed earlier, the assumption is that there are minimal changes in total temperature across the OGVs, so the fan exit total temperatures are used in the calculation of the stage adiabatic efficiency. The stage performance results for the fixed area nozzle operating lines for same tip clearance configurations are also shown in the figure. The differences in stage pressure ratio and adiabatic efficiency with changes in fan tip clearance are nearly the same compared with the differences shown in the fan performance maps (Fig. 9) for all speed lines and at all weight flows. A better method to visualize the stage performance differences between tip clearance configurations is shown in Figures 12 and 13. Here, the total pressure and adiabatic efficiency losses across the OGVs are expressed as percentage coefficients. Figure 12 shows the total pressure loss coefficient for the stage map and fixed operating lines for the nominal,.020 and.030 tip clearances. It can be seen that the pressure loss across the OGVs remain very small as the tip clearance increases. Larger differences can be seen at the far off design points located on either end of the bucket shaped speed line curves at all fan speeds. The performance differences between tip clearances become more significant at the higher fan speeds (95% and 100%) and at the weight flow extremes on each speed line. However, differences in total pressure loss are very small everywhere, only about 0.05% or less across the entire speed curve. Interestingly, the distinct grouping of the fixed nozzle operating lines into the two groups that was observed earlier in the fan performance plots can be seen in this figure. The nominal and.020 clearance results are in one group with nearly the same percent loss in pressure, and the.030 and.040 clearances in the other grouping with again nearly the same percent loss. The nominal/.020 group shows the lower pressure loss at all fan speeds. The difference in pressure loss between the two groups is small, about 0.05% at fan speeds above 77.5% corrected. In Figure 13 the loss in adiabatic efficiency across the OGVs for the map envelope is shown for the nominal,.020 and.030 tip clearances, and for the fixed nozzle operating line for all four tip clearances. As expected from the previous total pressure loss results, the adiabatic efficiency losses are very small at all fan speeds and weight flows. The largest performance differences with tip gap occur at the highest fan speeds, but are still only about 0.25% on the 100% corrected fan speed line. It would appear that the fan tip clearance has a minor effect on the OGV, or stage, performance. B. Farfield Acoustics The RAN configuration showed the most change in far field noise levels with changes in rotor tip clearance. Figure 14 shows 59 Hz bandwidth PWL spectra up to 20 khz for the RAN configuration at five rotor speeds. Blade 10

11 passage tones (nbpf) are not present at subsonic rotor tip speeds for the RAN configuration since there is no stator to interact with the rotor wake (61.7 and 75% corrected fan speed). However, there is clearly a broadband noise effect with changes in tip clearance at these speeds. At 61.7% corrected fan speed the broadband PWL is increased by as much as 2 db when the tip clearance increases from nominal to.030. However at 75% corrected fan speed and higher, the lowest broadband levels are seen for the.020 tip clearance rather than the nominal clearance. At 75% corrected fan speed the broadband noise level increases by as much as 1 db with the nominal and 2.5 db with the.030 clearance relative to baseline levels. This broadband noise increase is greatest at cutback, 87.5% corrected fan speed, where the noise level increase is 1.5 and 4 db for the nominal and.030 clearances, respectively, relative to the.020 clearance. At 90% corrected fan speed the broadband noise level increases about 3 db when the tip clearance is increased from.020 to.030 with a small increase from the baseline level for the nominal tip clearance. Broadband level changes associated with rotor tip clearance are negligible at 100% corrected fan design speed. The spectra provided in Figure 14 clearly show that the broadband noise level is most sensitive to changes in tip clearance at 87.5% corrected fan speed. LDV data presented previously (Ref. 12 and 15) show that: 1) a normal shock exists near 25% chord on the suction surface of the blade when the rotor is operating at this speed; 2) that this shock location is just downstream of where the tip clearance flow is generated; and 3) that the unsteadiness in the tip clearance flow decreases steadily with downstream distance. These data indicate, therefore, that the tip clearance flow passes through the normal shock when the tip flow is very unsteady (before it has had a chance to dissipate). This interaction between the shock and the unsteady tip flow may account for the increased broadband noise sensitivity which occurs when operating at this speed. These changes in the broadband noise level with tip clearance suggest that there is an optimum rotor tip clearance for minimum noise and that this optimum clearance is somewhat greater than the nominal, line-to-line clearance. Clearly, this broadband noise is being generated at the rotor tip, since there is no possibility for stator noise generation (although there is a possibility for turbulence scrubbing downstream of the rotor). At nominal clearance, the increase in broadband noise at nominal clearance relative to the.020 baseline clearance as the fan speed increases to 75% and above is most likely the result of a strongly blade tip interaction with the inlet boundary layer perhaps a scraping vortex that is formed ahead of the blade and interacts with the blade. The increase in broadband at.030 clearance relative to the.020 baseline may result from the larger/more turbulent tip clearance leakage flow interacting with the fan blade at the larger clearance. Rotor tones only appear for the RAN configuration at transonic and higher rotor tip speeds where these tones are generated by the shocks on the fan blades. Changes in rotor tip clearance had a minimal effect on rotor tones for the RAN configuration except for at 87.5% corrected fan speed where the BPF tone increased by 2 db as the tip clearance was increased from nominal to.030. At this speed the fan is transitioning to sonic conditions. The higher, more turbulent leakage flow through the more open tip clearance seems to be affecting the shock noise. Figure 15 shows the flyover EPNL for the RAN configuration with the three rotor tip clearances. Results are shown both as absolute EPNL and as delta EPNL relative to the.020 tip clearance, which was used as the baseline tip clearance for this test, as functions of rotor tip speed. Again, the noise level shows a regular increase with increasing tip clearance at the lowest tip speeds, whereas at higher subsonic through transonic rotor speeds the.020 tip clearance shows the lowest noise. These are the same noise trend with tip clearance seen in Figure 14. Figure 16 presents PWL spectra for the R4 rotor and the 54 vane, radial Baseline OGVs (cutoff with respect to the BPF tone at lower fan speeds). Changes in rotor tip clearance had very little impact on far field noise levels at any fan speed for this rotor/stator configuration. Flyover EPNL results are shown for this configuration in Figure 17, again showing that there is essentially no change in far field noise with changes in rotor tip clearance except for a small effect near transonic rotor tip speeds. At transonic tip speed (90% corrected fan speed), the noise level was slightly reduced with the nominal clearance and slightly increased with the.030 clearance relative to the baseline,.020 tip clearance. It appears that the rotor/stator interaction broadband noise created by the presence of the OGVs masks the changes in broadband noise with tip clearance, which were evident in the rotor alone data (Fig. 14 and 15). Far field noise results for the 26 vane, radial Low Count OGVs (which are cut-on) and the 26 vane, swept Low Noise OGVs (also cut-on) show somewhat more sensitivity to changes in rotor tip clearance, but not to the extent 11

12 seen for the RAN configuration. This is somewhat expected since previous SDT acoustic results (Ref. 13) have shown that as the rotor/stator interaction broadband noise decreases as the number of vanes decreases (54 to 26 radial) and as aft sweep is introduced with the fewer number of vanes (26 radial to 26 aft swept). Thus, as the rotor/stator interaction broadband noise decreases, the noise changes due to tip clearance are more evident. Figure 18 shows PWL spectra for the R4 rotor and Low Count OGVs. Broadband noise changes due to changes in tip clearance were only seen at transonic speeds, with the broadband noise about 1 db lower for the nominal tip clearance at 87.5% corrected fan speed. This result is difference than the RAN broadband noise results at the same fan speed (87.5% corrected fan speed) (Fig. 15), where the lowest noise levels were seen for the.020 tip clearance. The BPF tone level at 100% fan speed with the Low Count OGVs is about 2 db higher with the nominal tip clearance, and is significantly stronger than the corresponding BPF tone for the RAN configuration at this same fan speed. The fact that nbpf tones increase significantly when stators are present vs. rotor alone is noteworthy since it indicates that these tones are not due to shock noise - if they were shock noise related, they would show up in the RAN data; instead, these nbpf tones occurring at 100% speed must be due to rotor/stator interaction noise. Figure 19 shows the EPNL as a function of rotor tip speed for R4 and the Low Count OGVs. The EPNL trends with rotor tip clearance are somewhat similar to what was seen for the RAN configuration at subsonic rotor tip speeds (Fig. 15) in that the noise levels increase with increasing tip clearance. However, the magnitude of this change is much less than was seen in the RAN data. Once again, the result are indicating that most of the noise level changes associated with changes in tip clearance are masked by the additional broadband noise source created when stators are installed in the model i.e. rotor/stator interaction broadband noise. At transonic rotor tip speeds, there is a about a 1 EPNdB reduction associated with the nominal tip clearance relative to noise levels for the.020 clearance and even a slight noise reduction for the.030 clearance. These trends near transonic fan tip speeds are very much different than the RAN results, but again the noise levels are higher with the OGVs than without for RAN. The.020 tip clearance generated the highest EPNL near takeoff (100% corrected fan speed), with a significant drop for the nominal and.030 clearance (-1 to -1.6 db) at takeoff. The reason for this difference is not clear, but must be a result of the size and turbulence of the tip vortex and its interaction with the OGVs. Sound power level spectra for R4 with the swept Low Noise OGVs are shown in Figure 20. Noise changes associated with changes in fan tip clearance are typically small for this rotor/stator configuration. The rotor tone levels at 100% fan speed are again slightly higher with the nominal tip clearance. The flyover EPNL for the R4/Low Noise OGVs configuration in Figure 21 likewise show only small changes with changes in fan tip clearance. There is a 0.5 EPNdB noise reduction relative to that for the.020 tip clearance for the nominal tip clearance at subsonic tip speeds, while the larger.030 tip clearance shows no noise effect at lower subsonic tip speeds, and a 0.5 EPNdB reduction at higher subsonic tip speeds. There is little change in the EPNL at transonic and higher tip speeds associated with changes in the fan tip clearance. This is different than the results with the Low Count OGVs. Since the Low Noise vanes are swept aft compared to the Low Count vanes (one vane chord length at the vane tip), the increase in axial spacing is allowing the fan tip vortex to dissipate and mix out further before interacting with the vanes. As a result, while the trend in the noise data is similar compared to the Low Count OGV results, the difference is the noise levels is smaller, especially at the higher transonic and supersonic fan tip speeds. C. LDV surveys The LDV data were obtained in order to determine how the tip clearance flow downstream of the rotor changes as the tip gap changes. Prior to the test it was anticipated that increasing the tip gap at a given rotor speed would increase the flow disturbance created in the tip region and the overall level of noise produced by the turbofan model. The detailed measurements of the tip clearance flow made possible using LDV were expected to help explain these anticipated increases in noise. Figure 22 illustrates how the tip flow measured at the approach condition varies with tip gap. Parts a, b, c, and d of this figure show the tip flow as measured with the.040,.030,.020, and nominal (.000 ) rubstrips installed, respectively. Four different color contour plots are provided for each clearance: mean axial velocity in the upperleft; mean tangential velocity in the upper-right; axial turbulent velocity in the bottom-left; and tangential turbulent velocity in the bottom-right. Each contour plot shows the flow in the outer 25% of span at an axial location approximately 3.2 axial chord lengths (based on the tip) downstream of the tip trailing edge (Fig. 8). These data show very little change in the tip flow as the tip gap is varied at this low speed condition. 12

13 As rotor speed increases, the centrifugal force on the blades increases, the blades deform and untwist, essentially get longer, and the tip gap decreases. The centrifugal load on a blade operating at the approach condition (61.7% corrected fan speed) is much lower than at take-off (100% corrected fan speed). Consequently at the approach speed the tip gap associated with a given rubstrip is likely to be significantly larger than the value used to identify that rubstrip (i.e. the.040,.030, etc. value). These rubstrip identifying values correspond to the expected mean tip gap only when operating at 100% fan speed. At low speeds such as approach, a relatively large tip gap exists regardless of what rubstrip is installed. This may account for the insensitivity of the tip flow to changes in the tip clearance indicated by the data of Figure 22. Another reason for the relative lack of change in the tip clearance flow with tip gap in the Figure 22 data may be that much of the disturbance depicted in these plots is not really a tip clearance flow, per se. As pointed out in Reference 12, tip clearance flows get more pronounced relative to other viscous flows (boundary layers along the case and blades) as rotor speed and blade loading increases. At low speed the viscous flow associated with the tip leakage vortex can be buried in the relatively thick boundary layer flows which form on the case and blades. Therefore, even though the tip gap flow may change as the tip clearance changes it would be difficult to see those changes if the tip flow is buried in the boundary layer. Figure 23 shows the LDV data obtained at the cut-back operating condition (87.5% corrected fan speed). These data show more pronounced changes in the flow with changes in the tip gap than were evident in the approach condition data. These flow changes are especially evident in the tangential turbulent velocity contour plots. This is noteworthy because this quantity is expected to be important in regards to rotor/stator interaction broadband noise that is, changes in the tangential turbulent velocity are expected to lead to changes in the amount of broadband noise generated when this flow interacts with the downstream stator vanes. The Figure 23 data show a significant reduction in both the peak level of tangential turbulent velocity and the size of the turbulent region associated with the tip gap flow as the tip clearance is decreased. Based on this, one could expect a reduction in the amount of broadband noise produced when the tip clearance flow interacts with the OGVs. The acoustic spectra of Figures 16, 18 and 20 show a decrease in the broadband level, and the data of Figures 17, 19 and 21 show an overall drop in EPNL of about 1 db when the tip clearance is reduced from.030 to nominal. These changes, however, are surprisingly small considering the rather drastic change in the tip clearance flow depicted in Figure 23. Apparently, this noise source the tip clearance flow - is masked by another, more dominant source perhaps blade and/or vane self-noise or blade-wake/stator interaction noise. Strong supporting evidence in favor of the argument that the blade wake/stator interaction noise is masking any noise reduction benefit in reducing the tip gap is shown in Figures 14 and 15. Here, the acoustic spectra and EPNL plots for the isolated R4 rotor, available using the RAN model hardware, are shown. It can be seen that reducing the tip gap produces larger broadband noise reductions as well as larger reductions in EPNL at all fan speeds compared to the acoustic results with OGVs installed. However, evidence of another tip flow noise generating mechanism can also be seen since there is a crossover point between the nominal clearance and the.020 clearance where the nominal tip clearance is generating more noise. This suggests that there may indeed be an increased interaction of the fan tip flow with the inlet boundary layer for the nominal tip gap. In addition, the size of the tip flow vortex and the way it interacts with the 26 vane cut-on OGVs compared with the 54 vane cut-off OGVs suggests that a noise benefit is achieved around this fan speed since there is a reduction in EPNL noise of about 0.5 db demonstrated for the nominal and.030 tip gaps as shown in Figures 19 and 21. Finally, the drop in blade loading at the fan tip can be seen in the mean axial velocity plots in this Figure 23 as the strength and depth of the wake deficit decreases with the increases in tip clearance. Also, the mean tangential velocity plots show the development of the leakage flow and growth of the tip vortex as the tip clearance increases. Figure 24 shows data acquired at the highest speed at which LDV data were acquired, the take-off condition (100% corrected fan speed). At this speed the plots of all four measured quantities show significant changes in the tip flow as the tip gap changes. Based on these changes in the tip clearance flow one might expect significant reductions in the noise generated by the model as tip gap is reduced. The EPNL plots of Figure 17, 19 and 21 indicate, however, that these noise reductions did not occur. Once again, this noise source appears to be masked by one or more other sources perhaps a combination of shock-induced multiple pure tone and rotor-wake/stator vane interaction noise. 13

14 VI. Summary and Conclusions The design of effective new technologies to reduce aircraft propulsion noise is dependent on identifying and understanding the noise sources and noise generation mechanisms in the modern turbofan engine, as well as determining their contribution to the overall aircraft noise signature. Therefore, a comprehensive aeroacoustic wind tunnel test program was conducted as part of the NASA Quiet Aircraft Technology program called the Fan Broadband Source Diagnostic Test. The test was performed in the anechoic NASA Glenn 9- by 15-Foot Low Speed Wind Tunnel using a 1/5 scale model turbofan simulator that is representative of a current generation, medium pressure ratio high bypass turbofan engine. The investigation focused on the simulated bypass section of the turbofan engine. The technical objectives of the test were: 1) to identify the noise sources within the model and their contribution to the overall noise level; 2) to investigate several component design technologies by evaluating their impact on the aerodynamic and acoustic performance; and 3) to conduct detailed flow diagnostics within the research model to help in understanding the physics of the flowfield. Details were presented in this report on the effect of the bypass fan tip clearance on the bypass stage fan and outlet guide vanes (OGVs) aerodynamic performance and farfield acoustics, in addition to fan wake flow diagnostic results. Four different fan design point tip clearances were investigated in combination with a 22-blade, 22-inch diameter baseline fan (R4) and a 54-vane, radial, acoustically cut-on set of OGVs (Baseline OGVs). The four fan tip clearances were: nominal, or line-to-line (0.000 ), representing a new engine;.020, represent a typical older engine about to be removed from service for maintenance, or at end of life;.030, representing the clearance after the aircraft experiences hard maneuvering or a hard landing; and.040, representing the clearance after a severe aircraft landing. Test results on the effect of fan tip clearance on the farfield acoustics for three additional OGV configurations rotor alone (no OGVs), 26 vane radial, and 26 vane swept were also discussed. The aerodynamic results obtained showed: 1. There are small changes in fan performance with changes in fan tip clearance at fan speeds below 87.5% on the pressure ratio, temperature ratio and adiabatic efficiency performance maps. Above that speed, the changes are small. On the 100% corrected fan design speed line, the loss in peak efficiency is about 0.5% from nominal to.020 clearance (from 93.6% to 93.1%), and 1.1% from nominal to 0.30 clearance (from 93.6% to 92.5%). The drop in performance especially at the higher fan speeds is most likely related to the leakage flow around the tip gap which caused a drop in the blade loading locally. 2. The fan performance on the fixed area nozzle operating lines show that differences in fan performance only become significant as the tip clearance increases beyond.020 at all fan speeds, but especially above 77.5%. At fan speeds below 77.5%, there was only a 0.5% maximum change from nominal to.040 clearance. As the fan speed increased to 87.5%, the.030 and.040 clearance data showed the efficiency start to fall off compared to the nominal and.020 clearances. Above 95% speed, more significant differences occur at any clearance larger than the nominal. Also, beyond.030 clearance it appears that there are minor changes in performance due to the additional clearance, since the data showed that the fan performance for the.030 and.040 clearance configurations were very similar. This indicated that no further drop in blade loading at the tip occurred and the tip leakage flow had essentially stabilized beyond.030 clearance. At 100% corrected fan speed, where the largest differences in performance occurred, the.030 and.040 clearance differences are still small (less than 0.25%), but the.020 clearance performance falls off more than the nominal clearance. At this speed, the efficiency for the nominal clearance was 91.9%, the.020 was 91.5%, the.030 was 91.1% and the.040 was 90.9%. The difference in corrected weight flow for this speed ranges from 97.2 lbs/sec to 97.4 lbs/sec from.040 to nominal clearance, respectively. 3. Fan performance profiles verified that the drop in performance as tip clearance increases resulted from lower performance at the most outboard portion of the fan blade - beyond 90% of blade span for pressure ratio and 84% span for temperature ratio. The total temperature seems to be more sensitive than total pressure to changes in fan tip clearance. 14

15 4. Stage performance results showed that the OGVs are fairly insensitive to any changes in fan tip clearance below 87.5% fan speed. Even above that speed changes in adiabatic efficiency were very small, on the order of 0.25% from nominal to.040 clearance at 100% fan speed. The farfield acoustic results showed: 1. Changes in noise level due to changes in fan tip clearance are primarily generated by the fan at the fan tip region. 2. The rotor-alone (RAN) configuration provides the greatest insight into this mechanism, showing that broadband noise levels are typically lowest for the nominal clearance and highest at the.030 clearance at fan speeds below 75%. Above that speed, the 020 tip clearance shows the lowest broadband noise level and both the nominal and.030 clearances show higher noise, with the.030 being the loudest. A possible explanation is that at lower speeds the amount of the tip leakage flow and its interaction with the fan tip are driving the noise. Above 75%, the nominal spacing the rotor tip is adversely interacting with the inlet boundary layer causing a scraping vortex to be formed that interacts with the fan tip generating more broadband noise than the.020 clearance. As the tip clearance gets larger, the higher tip leakage flow vortex formed is driving the noise, and the larger the gap the higher the noise level. 3. The downstream OGVs, to a large degree, mask the farfield propagation of these rotor tip clearance induced noise changes since the rotor/stator interaction and broadband noise levels are higher than the rotor alone noise levels. However, the OGV vane number configuration had some effect on the amount of noise reduction with fan speed. The 54-vane radial, cut-off Baseline OGVs seemed to show that the noise was relatively unaffected by tip clearance except near transonic fan tip speeds. However, the 26-vane radial, cut-on Low Count OGVs and 26-vane swept, cut-on Low Noise OGVs showed trends in the noise differences with fan tip clearance more in line with the RAN results below 70% fan speed. Above 70% fan speed there was a distinct difference in the noise generated by tip clearance flow and rotor/stator interaction as the tip gap changed. Results showed the.020 was noisier than the other two clearances until the fan speed was transonic, anywhere from 0.5 to 1.0 EPNdB. 4. There appears to be a noise generation mechanism as the fan tip interacts with the wall boundary layer at small tip clearances, and as the fan tip leakage flow interacts with the fan blade tip at larger tip clearances. The rotor/stator interaction noise levels, both tone and broadband, also appear to be affected by the larger tip gap and leakage flow vortex that is generated. The rotor/stator interaction tone level at 100% fan speed may be slightly higher with the nominal tip clearance compared with the.020 and.030 clearances resulting from the fan/boundary layer interaction, possibly as a result of a scraping vortex being generated with the close proximity of the fan tip to the fan case. The LDV diagnostic results of the fan tip flow field showed: 1. There were minimal changes in the fan tip flow as the tip gap increased from nominal to.040 at the 61.7% approach fan speed. 2. At 87.5% cutback fan speed, more pronounced changes in were seen, especially in the tangential component of turbulence. A distinct increase in the size of the vortex flow could be seen as the tip gap increased. The tangential turbulence in the fan wake is thought to be the main contributing mechanism to the rotor/stator interaction broadband noise generated. However, the large changes in turbulence did not contribute significantly to broadband noise indicted in the farfield acoustic results. 3. At 100% takeoff fan speed, significant changes in the two components of both mean velocity and turbulence showed dramatic changes, which would indicate dramatic changes in the rotor/stator interaction noise. However, since the acoustics showed no appreciable change with tip gap, the assumption is that the fan blade/flow shock noise generated at the supersonic fan tip speeds is masking the change in acoustics that occur with the change in tip flow as the tip gap increases. 15

16 References 1 Groeneweg, J.F., Rice, E.J., Aircraft Turbofan Noise", Vol.109, Journal of Turbomachinery; January Groeneweg, J.F., Fan Noise Research at NASA, NASA TM , Envia, E., Fan Noise Reduction: An Overview, NASA/TM , AIAA , Bridges, J., Envia, E., Huff, D., Recent Developments in U.S. Engine Noise Reduction Research, NASA/TM , ISABE , Dittmar, J.H., Elliott, D.M., Bock, L.A., Some Acoustic Results from the Pratt and Whitney Advanced Ducted Propulsor - Fan 1 ; NASA/TM ; March Dittmar, J.H., Hughes, C.E., Bock, L.A., Hall, D.G., Cruise Noise Measurements Of A Scale Model Advanced Ducted Propulsor ; 15 th AIAA Aeroacoustics Conference, Long Beach, CA; October 25-27, Envia, E., Nallasamy, M., Design Selection And Analysis Of A Swept And Leaned Stator Concept ; NASA/TM , December Hoyniak, D., Fleeter, S., Effect Of Aerodynamic Detuning On Supersonic Rotor Discrete Frequency Noise Generation ; NOISE-CON 88, Purdue University, West Lafayette, IN; June 20-22, Woodward, R.P., Elliott, D.M., Hughes, C.E., Berton, J.J., "Benefits Of Swept And Leaned Stators For Fan Noise Reduction"; NASA/TM ; AIAA ; November Podboy, G.G., "Further Analysis of Fan Wake Data Obtained Downstream of the Allison Low Noise Fan and the Pratt & Whitney ADP Fan 1 Models," AST Engine Noise Workshop, Vol. IV, pp , April Hughes, C.E., Aerodynamic Performance of Scale-Model Turbofan Outlet Guide Vanes Designed for Low Noise, AIAA , January Podboy, G.G., Krupar, M.J., Helland, S.M., Hughes, C.E., Steady and Unsteady Flow Field Measurements within a NASA 22 Inch Fan Model, AIAA , January 2002/ NASA TM , Woodward, R.P., Hughes, C.E., Jeracki, R.J., Miller, C.J., Fan Noise Source Diagnostic Test Far-field Acoustic Results, AIAA , June, Hughes, C.E., Jeracki, R.J., Miller, C.J., Fan Noise Source Diagnostic Test Rotor Alone Aerodynamic Performance Results, AIAA , June Podboy, G.G., Krupar, M.J., Hughes, C.E., Woodward, R.P., Fan Source Diagnostic Test LDV Measured Flow Field Results, AIAA , June Heidelberg, L.J., Fan Noise Source Diagnostic Test Tone Model Structure Results, AIAA , June Yuska, J.A., Diedrich, J.H., Nestor, C., Lewis 9- by- 15-Foot V/STOL Wind Tunnel, NASA TM X 2305, Arrington, A.E., Gonsalez, J.C., Flow Quality Improvements in the NASA Lewis Research Center 9- by 15-Foot Low Speed Wind Tunnel, NASA CR , Dahl, M.D., 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, April Dahl, M.D., Woodward, R.P., Background Noise Levels Measured I the NASA Lewis 9- by 15-Foot Low Speed Wind Tunnel, NASA TP-3274, November Woodward, R.P., Dittmar, J.H., Background Noise Levels Measured in the NASA Lewis 9- by 15-Foot Low-Speed Wind Tunnel, NASA TM , AIAA , January Balan, C., Hoff, G.E., Propulsion Simulator for High Bypass Turbofan Performance Evaluation, SAE , Jeracki, R.J., Model Engine Performance Measurement from Force Balance Instrumentation, NASA TM /AIAA , July, Chestnutt, D., Flight Effects of Fan Noise, NASA CP-2242, January,

17 Table 1 - Fan Design Parameters No. of Blades 22 Tip Diameter, in. 22 Inlet Radius Ratio 0.30 Corrected Design Speed, RPM 12,657 Design Tip Speed, ft/s 1,215 Corrected Weight Flow, lb m /s Specific Flow, lb m /s-ft Stage Pressure Ratio 1.47 Table 2 - Summary of fan and OGV airfoil geometries Span Fan Baseline Low Count Low Noise No. Blades/Vanes Aft Sweep, deg Aspect Ratio Pitchline Chord *, in Pitchline Solidity * Hub Pitchline Tip Stagger *,1,2, deg Hub Pitchline Tip Vane Camber *, deg Hub Pitchline Tip t max /c * Hub Pitchline Tip * Pitchline values 1 Defined from axial plane; positive angle in direction of fan rotation. 2 Positive angle in opposite direction of fan rotation for OGVs. 17

18 Traversing Microphone Flow Fixed Microphones Test Section Centerline Fan Module UHB Drive Rig Fan Stacking Axis Model STA Figure 1. Top view schematic showing the location of the UHB Drive Rig in the 9- by 15-Foot Low Speed Wind Tunnel. Fan Module Location Air Turbine Drive Operating Capabilities Maximum Power 5,050 HP Maximum Speed 16,850 RPM Turbine Plenum Max Operation Pressure 230 psia Temperature 500 F Mass Flow 33.6 lbs/sec Drive Rig Support Strut Figure 2. NASA Glenn Research Center Ultra High Bypass (UHB) Drive Rig propulsion simulator. Flight-Type Nacelle Fixed Area Flight-Type Nozzle Flight-Type Inlet Fan Baseline OGVs Static Force Balance UHB Drive Rig Rotating Balance Figure 3. Schematic diagram of the Fan Model and Baseline OGVs installed on the UHB Drive Rig. 18

19 54 Radial 26 Radial 26 Swept Figure 4. Schematic illustrating orientation of OGV configuration in the fan module. Variable area Fan Exit Nozzle Fan Case/Rubstrip Bellmouth Inlet a). Fan module with bellmouth inlet and variable area fan exit nozzle. b). Close-up view of variable fan exit nozzle hardware arrangement. Figure 5. Photographs of fan module installed in the NASA Glenn 9- by 15-Foot Low Speed Wind Tunnel in the aerodynamic performance configuration. 19

20 UHB Drive Rig air turbine Air supply and support strut a). Fan module with flight-type inlet, nacelle and fixed area fan nozzle. Fixed microphones Traversing microphone b). View of acoustic microphone orientation in the wind tunnel. Figure 6. Photographs of fan module installed in the Glenn 9- by 15-Foot Low Speed Wind Tunnel in the farfield acoustic configuration (concluded). 20

21 Cylindrical shield covering optics a). LDV system hardware installed next to the research fan model in the wind tunnel. receiving optics windows fiber optic cables transmitting optics b). LDV system components with optics protective shield removed. Figure 7. LDV system components and installation setup with the fan module in the Glenn 9- by 15-Foot Low Speed Wind Tunnel (concluded). 21

22 Bellmouth Inlet Location of LDV Measurements Swept OGV Fan Flow Figure 8. Schematic illustrating location of LDV measurements in the fan module. 110 Corrected Weight Flow, ω f, lb m /sec Tip Clearance Nom X Corrected Fan Speed, % a). Corrected weight flow. Figure 9. Fan aerodynamic performance maps and on individual fixed nozzle operating lines (continued). 22

23 X Tip Clearance Nom % Fan Pressure Ratio, P t,f /P t, % 87.5% 95% 70% % 50% Corrected Weight Flow, ω f,c, lb m /sec b). Fan pressure ratio X Tip Clearance Nom % Fan Temperature Ratio, T t,f /T t, % 77.5% 87.5% 95% % 60% Corrected Weight Flow, ω f,c, lb m /sec c). Fan temperature ratio. Figure 9. Fan aerodynamic performance maps and on individual fixed nozzle operating lines (continued). 23

24 100 Fan Adiabatic Efficiency, ηf, % X Tip Clearance Nom % 60% 70% 77.5% 87.5% 95% 100% Corrected Weight Flow, ω f,c, lb m /sec d). Fan adiabatic efficiency. Figure 9. Fan aerodynamic performance maps and on individual fixed nozzle operating lines (concluded). Tip Tip Clearance Nom X Blade Span, % Hub % 75% Fan Rake Pressure Ratio, P t,r /P t,o a). Fan total pressure ratio. Figure 10. Fan STA 12.5 radial performance profiles from fixed rakes (continued) % 100%

25 Tip Tip Clearance Nom X Blade Span, % % 75% 87.5% 100% Hub Fan Rake Temperature Ratio, T t,r /T t,o b). Fan total temperature ratio. Tip Blade Span, % Tip Clearance Nom X Fan Speed 61.7% (AP) 75% 87.5% (CB) 100% (TO) Hub Fan Rake Adiabatic Efficiency, η f, % c). Fan adiabatic efficiency. Figure 10. Fan STA 12.5 radial performance profiles from fixed rakes (concluded). 25

26 X Tip Clearance Nom Stage Pressure Ratio, P t,s /P t, % 87.5% 95% 100% % 70% 50% Corrected Weight Flow, ω f,c, lb m /sec a). Stage pressure ratio X Tip Clearance Nom Stage Adiabatic Efficiency, ηs, % % 60% 70% 77.5% 87.5% 95% 100% Corrected Weight Flow, ω f,c, lb m /sec b). Stage adiabatic efficiency. Figure 11. Stage adiabatic efficiency map and on individual fixed nozzle operating lines (concluded). 26

27 2.6 Stage Total Pressure Loss Coefficient, (P t,f -P t,s )/P t,f, % X Tip Clearance Nom % 60% 70% 77.5% 87.5% 95% 100% Corrected Weight Flow, ω f,c, lb m /sec Figure 12. Stage total pressure loss across the OGVs on the stage operating map and on fixed operating lines. 13 Stage Adiabatic Efficiency Loss, (ηf - ηs), % X Tip Clearance Nom % 60% 70% 77.5% 87.5% 95% 100% Corrected Weight Flow, ω f,c, lb m /sec Figure 13. Stage adiabatic efficiency loss across the OGVs on stage operating map and on fixed operating lines 27

28 Rotor Tip Clearance Nominal (rub-in) " baseline 0.030" % speed (approach) Sound Power Level, db % speed BPF 2BPF 3BPF 4BPF % speed (cutback) % speed % speed (takeoff) Frequency, Hz. Figure 14. Sound power level spectra for the RAN configuration at five rotor speeds. 100 EPNL, db Delta 1 Flyover 0.5 EPNL, db Approach Cutback Takeoff Rotor tip clearance Nominal (rub-in) 0.020" baseline 0.030" Tangential tip speed, ft./s Figure 15. Flyover EPNL for the RAN configuration. 28

29 Rotor Tip Clearance Nominal (rub-in) " baseline 0.030" % speed (approach) Sound Power Level, db % speed BPF 87.5% speed (cutback) 2BPF 3BPF 4BPF % speed % speed (takeoff) Frequency, Hz Figure 16. Sound power level spectra for the R4 fan and Baseline OGVs at five rotor speeds. EPNL, db Delta 1 Flyover 0.5 EPNL, db Approach Cutback Takeoff Rotor tip clearance Nominal (rub-in) 0.020" baseline 0.030" Tangential tip speed, ft./s Figure 17. Flyover EPNL for the R4 fan and Baseline OGVs. 29

30 % speed (approach) 125 Rotor Tip Clearance Nominal (rub-in) " baseline 0.030" Sound Power Level, db % speed BPF 2BPF 3BPF 4BPF % speed (cutback) % speed % speed (takeoff) Frequency, Hz Figure 18. Sound power level spectra for the R4 fan and Low Count OGVs at five fan speeds 100 Cutback EPNL, db Delta Flyover EPNL, db Approach Takeoff Rotor tip clearance Nominal (rub-in) 0.020" baseline 0.030" Tangential tip speed, ft./s Figure 19. Flyover EPNL for the R4 fan and Low Count OGVs. 30

31 Rotor Tip Clearance Nominal (rub-in) " baseline 0.030" % speed (approach) Sound Power Level, db % speed BPF 87.5% speed (cutback) 2BPF 3BPF 4BPF % speed % speed (takeoff) Frequency, Hz Figure 20. Sound power level spectra for the R4 fan and Low Noise OGVs at five rotor speeds EPNL, db Delta Flyover EPNL, db Approach Cutback Takeoff Rotor tip clearance Nominal (rub-in) 0.020" baseline 0.030" Tangential tip speed, ft./s Figure 21. Flyover EPNL for the R4 fan and Low Noise OGVs. 31

32 Direction of Rotation All views Aft Looking Forward Mean Axial Velocity Mean Tangential Velocity Mean Axial Velocity Mean Tangential Velocity Axial Turbulent Velocity Tangential Turbulent Velocity Axial Turbulent Velocity Tangential Turbulent Velocity a)..040 rubstrip b)..030 rubstrip Figure 22. LDV data measured at 61.7% corrected fan speed (approach condition) (all velocities in ft/sec) (continued). 32

33 Direction of Rotation All views Aft Looking Forward Mean Axial Velocity Mean Tangential Velocity Mean Axial Velocity Mean Tangential Velocity Axial Turbulent Velocity Tangential Turbulent Velocity Axial Turbulent Velocity Tangential Turbulent Velocity c)..020 rubstrip d). Nominal rubstrip Figure 22. LDV data measured at 61.7% corrected fan speed (approach condition) (all velocities in ft/sec) (concluded). 33