Robert J. Jeracki

Glenn Research Center, Cleveland, Ohio

Comprehensive Report of Fan Performance

From Duct Rake Instrumentation on

1.294 Pressure Ratio, 806 ft/sec Tip Speed

Turbofan Simulator Models

NASA/TM—2006-213863

February 2006
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Robert J. Jeracki

Glenn Research Center, Cleveland, Ohio

Comprehensive Report of Fan Performance

From Duct Rake Instrumentation on

1.294 Pressure Ratio, 806 ft/sec Tip Speed

Turbofan Simulator Models

NASA/TM—2006-213863

February 2006

National Aeronautics and

Space Administration

Glenn Research Center
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NASA/TM—2006-213863 iii

Contents 

Page 

Contents............................................................................................................................. iii 

Abstract............................................................................................................................ 1 

Nomenclature.................................................................................................................... 1 

Introduction....................................................................................................................... 2 

Fan Performance Test Hardware...................................................................................... 3 

Instrumentation for Fan Performance................................................................................ 3 

Fan Performance Parameter Calculations......................................................................... 4 

Fan Blade Configurations.................................................................................................. 6 

Fan Operating Line Results and Repeatability, TI Blade, –9 Degrees, FCT Rubstrip......... 6 

TI Versus GC Fan Operating Line Results at –9 Degree Angle, Smooth Rubstrip.................. 7 

TI Versus GC Fan Operating Line Results at –15 Degree Angle, Smooth Rubstrip................ 8 

–9 and –15 Degree Angle Fan Operating Line Results, TI Blade, Smooth Rubstrip................ 9 

–9 and –15 Degree Angle Fan Operating Line Results, GC Blade, Smooth Rubstrip............. 9 

Smooth Versus FCT Rubstrip Operating Line Results, TI Blade, –9 Degree Angle................ 10 

Fan Performance Map Results and Repeatability, GC Blade, –9 Degrees, 

Smooth Rubstrip.................................................................................................................. 11 

TI Versus GC Fan Map Results at –9 Degree Angle, Smooth Rubstrip................................... 12 

TI Versus GC Fan Map Results at –9 Degree Angle, FCT Rubstrip........................................ 13 

Smooth Versus FCT Rubstrip Fan Map Results, GC Blade, –9 Degree Angle........................ 13 

Smooth Versus FCT Rubstrip Fan Map Results, TI Blade, –9 Degree Angle.......................... 14 

Smooth Versus FCT Rubstrip Fan Map Results, TI Blade, –15 Degree Angle........................ 15 

Smooth Versus FCT Rubstrip Fan Map Results, TI Blade, 0 Degree Angle........................... 15 

TI Versus GC Single-Radius Map Results at –9 Degree Angle, Smooth Rubstrip................... 16 

TI Versus GC Single-Radius Map Results at –9 Degree Angle, FCT Rubstrip........................ 17 

Smooth and FCT Single-Radius Map Results, TI Blades, –9 Degree Angle........................... 18 

TI Blade Instability Occurrences........................................................................................ 19 

Summary of Results.......................................................................................................... 20

References........................................................................................................................ 22 

NASA/TM—2006-213863 1

Comprehensive Report of Fan Performance from Duct Rake 

Instrumentation on 1.294 Pressure Ratio, 806 ft/sec 

Tip Speed Turbofan Simulator Models 

Robert J. Jeracki 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

A large scale model representative of an advanced ducted propulsor-type, low-noise, very high 

bypass ratio turbofan engine was tested for acoustics, aerodynamic performance, and off-design 

operability in the NASA Glenn 9- by 15-Foot Low-Speed Wind Tunnel. This test was part of 

NASA’s Advanced Subsonic Technology Noise Reduction Program. The low tip speed fan, nacelle, 

and an un-powered core passage (with core inlet guide vanes) were simulated. To measure fan 

aerodynamic performance, four rakes with total pressure and temperature probes were installed in 

the bypass duct, upstream of the stators. Core inlet guide vanes were instrumented with total 

pressure or total temperature probes, measuring performance in the fan hub region. Several wind 

tunnel runs with nearly the same hardware installed are compared, to identify the repeatability of 

the fan performance measurements. In other wind tunnel runs, fan performance was measured 

with two different fan blades made from different material and intended for the same design 

operation (graphite composite material over a titanium spar and solid titanium), three fan blade 

angles (cruise design, –9 degrees lower for takeoff, and even lower at –15 degrees), and two fan 

rubstrip designs (smooth and a casing treatment for stall management). 

As might be expected, the effect of the stall management casing treatment was a performance 

penalty. The amount of recirculating flow at the fan tip was larger than required for the desired stall 

margin improvement, however. Reducing that flow would also reduce the resulting performance 

penalty. There was a somewhat surprising performance difference between the titanium and 

graphite composite blades, mostly away from the fan operating speed lines. Though both blades 

showed some indication of a flow transition (near the blades tip) as the fan flow was reduced, the 

transition occurred near the takeoff operating lines with the titanium blades and resulted in high 

blade stress instabilities. 

Though the pressure and temperature ratios were low for this fan design, the techniques used 

to improve thermocouple measurement accuracy gave repeated data runs with the adiabatic 

efficiency agreeing within one percent. The measured fan adiabatic efficiency at takeoff was 

93.7 percent which matched the design intent well. 

Nomenclature 

AARPT12.5 Area average of all 10 total pressure averages in the bypass duct, psi 

AARTT12.5 Area average of all 10 total temperature averages in the bypass duct, °R 

AARPTCI Area average of all 5 total pressure averages in the core inlet, psi 

AARTTCI Area average of all 5 total temperature averages in the core inlet, °R 

ADP Advanced Ducted Propulsor 

ARPT12.5 STA 12.5 total pressure averaged at each of 10 bypass duct radii, psi 
NASA/TM—2006-213863 2

ARTT12.5 STA 12.5 total temperature averaged at each of 10 bypass duct radii, °R 

ARPTCI Core inlet total pressure averaged at each core inlet radius, psi 

ARTTCI Core inlet total temperature averaged at each core inlet radius, °R 

CFD Computational Fluid Dynamic (computer flow calculations) 

EADAAR12 Bypass adiabatic efficiency, (RAARPT12(2/7)-1)/(RAARTT12–1) 

FCT Fan Casing Treatment, rubstrip design for stall margin improvement 

GC Graphite Composite blade material 

IGV Inlet Guide Vanes of the core passage 

Mo Freestream Mach number 

Mlocal Mach number based on specific local pressure measurements 

MS Axial model station referenced to the fan stacking axis as 170.000, in 

PSTD Standard day pressure, 14.696 psi 

PS Static pressure, psi 

PT Total pressure, psi 

PTo Freestream total pressure, psi 

RAARPT12 Bypass pressure ratio, AARPT12.5/PTo 

RAARTT12 Bypass temperature ratio, AARTT12.5/TTo 

RPMc Rotor rpm corrected to standard day conditions, rpm 

RPMck RPMc/1000, rpm/1000 

STA Station, in standard engine nomenclature (i.e., 0 is far upstream, 2.0 is just 

upstream of the fan, 12.5 is between the fan and bypass duct stator, ∞ is far 

downstream) 

TSTD Standard day temperature, 518.67 °R 

TI Titanium blade material 

TPS Turbine Powered Simulator 

TTo Freestream total temperature, °R 

WBc Bypass weight flow corrected to standard day conditions, lbm/sec 

WCc Core weight flow corrected to standard day conditions, lbm/sec 

WCc,loc Core weight flow corrected to local core conditions, lbm/sec 

WFc Fan total weight flow corrected to standard day conditions, lbm/sec 

Introduction 

Fan engine performance test methods have been developed, refined, and standardized 

through years of experience (refs. 1 and 2). Methods include component tests (inlet, fan, stage, 

and nozzle), ground test, and altitude test facilities. Flow-through nacelles and small Turbine 

Powered Simulator (TPS) models have normally been used to evaluate installation effects. Larger 

TPS models with strain gaged force balances may be used to measure fan engine thrust-minusdrag performance in wind tunnels (ref. 3). Total pressure and total temperature rakes can be 

installed to measure fan and/or stage performance. NASA Glenn is continuing to test 22-inch 

diameter fan models of existing and advanced engines to evaluate low noise improvement 

possibilities and operability or performance improvements (refs. 4 to 8). This is part of NASA’s 

Aeronautics and Space Transportation Technology Enterprise. This report presents fan 

performance results of an ADP-type, low tip speed, low pressure ratio fan, using fan duct rake total 

pressure and temperature measurements. 

A 22-inch fan diameter model simulating a low tip speed, Advanced Ducted Propulsor 

(ADP)-type engine is shown in the NASA Glenn 9- by 15-Foot Low-Speed Wind Tunnel (fig. 1(a)), 

(ref. 9). The tunnel can be run from near zero to 0.2 Mach number, at near ambient barometric 

pressure, and temperatures from 70 to 75 °F. The simulator is mounted to the front of a strut-
NASA/TM—2006-213863 3

supported high-pressure air turbine. The strut carries heated air to the turbine, lubrication lines to 

and from the model, and instrumentation leads off of the model to the high-speed data system. A 

cross section of the simulator is shown in figure 1(b). The engine core simulation consists of an 

inlet, inlet guide vanes, a flow-through duct with support struts, and a nozzle. The core nozzle exit 

area was expanded to pass the correct core flow at takeoff conditions, so it does not represent the 

exact geometry of the engine being simulated. If the core were modeled with a first compressor 

stage, the core nozzle simulation could match the engine core nozzle geometry more closely. 

Fan Performance Test Hardware 

The fan engine simulator included a flight-type inlet and nozzle. Some fan performance testing 

with the duct pressure and temperature probes was done with the flight hardware installed. The results 

are along a specific operating line of pressure ratio versus fan flow for each freestream Mach number 

tested. To obtain fan performance maps, a bellmouth inlet and a Variable Fan Exit Nozzle (VFEN) 

were installed. The bellmouth allowed the fan to be run at low freestream Mach number without inlet 

flow separations or ingested vortices. The VFEN allowed the nozzle exit area to be reduced, with 

remotely actuated radial blades acting like butterfly valves. Reducing the nozzle area made the fan run 

at lower mass flow, through a range of engine operating conditions from high-flow simulating cruise, to 

lower flow for takeoff, and finally to fan stall. This report is organized to present flight nozzle operating 

line performance comparisons first, followed by fan performance maps 

Instrumentation for Fan Performance 

Freestream total pressure and temperature were measured using a cruciform-shaped rake 

assembly located about 86.5 inches in front of the fan blade stacking axis. There were a total of 

four pressure probes and five temperature probes located along the arms of the cruciform and 

centered on the fan centerline axis, accurately measuring the flow conditions coming into the fan. 

The thermocouple probes were shielded with aspirated stagnation tubes. For acoustic data runs, 

the freestream cruciform rake and the bypass duct rakes were removed. 

Axial locations of instrumentation in the engine simulator are shown in figure 2(a). (Note: MS 

stands for Model Station in inches, referenced to the fan blade stacking axis as 170.000.) Fan total 

weight flow was calculated upstream of the fan (standard engine nomenclature Station 2.0, ref. 10) 

using a ring of nine static pressures in the inlet at MS 164.440 (figs. 2(a) and 3(a)). Core weight 

flow was calculated using two rakes of five total pressures each, and two rings of four static 

pressures each, located at MS 195.850, upstream of the core nozzle exit (figs. 2(a) and 3(b)). 

Fan performance was measured using four radial rakes midway through the stage bypass duct 

(engine STA 12.5) at MS 175.000 (figs. 2(a) and 4(a)), and probes in the core inlet (MS 173.844–

173.946) attached to the core Inlet Guide Vanes (IGVs) (figs. 2(a) and 4(b)). The four bypass duct 

rakes had pressure and temperature probe paired together at 10 radial locations, at the centers of 

equal annular area (a total of 40 pressures and 40 temperatures). The core inlet pressures and 

temperature probes were located at the centers of 5 equal area annuli, with two measurements at 

each radius (a total of 10 pressures and 10 temperatures). The bypass duct pressure tubes and 

thermocouples were shielded with aspirated stagnation tubes, giving high flow angle capability. 

The thermocouple shields had very high recovery ratio and, therefore, small temperature correction 

due to flow velocity. The core pressure tubes and thermocouples needed to be smaller, so the 

shields were smaller and the temperature recovery corrections were larger than the bypass duct 

probes. 

For some of the later runs, the core inlet pressure and temperature probes were missing. They 
NASA/TM—2006-213863 4

had been removed to eliminate any possible acoustic contamination during the far-field acoustic 

testing. For those later test runs, only bypass duct performance is available. 

Figure 2(b) is a photograph that shows the fan with the inlet and some of the fan blades 

removed. Three of the four bypass duct rakes can be seen in front of the stators. Some core 

probes are also somewhat visible on the core IGVs. 

Fan Performance Parameter Calculations 

For the bypass duct total pressures, the four total pressures from each duct rake were 

averaged at each of the ten annular radii in the bypass (ARPT12.5 (1 to 10)). At each rake radius, 

a linear interpolation was done from the averaged inner-wall static pressure to the averaged outerwall static pressure. Each static pressure and the corresponding averaged total pressure were 

used to calculate a local Mach number. A total temperature recovery ratio had been measured in a 

calibration facility and correlated with Mach number. For the wind tunnel data, the indicated 

temperature of each thermocouple was corrected by the temperature recovery ratio corresponding 

to the calculated local Mach number at each radius. The four total temperatures from each duct 

rake were averaged at each of the ten annular radii in the bypass (ARTT12.5 (1 to 10)). Then, the 

ten annular averages of total pressure and temperature were averaged (based on equal areas) to 

yield a single value of averaged radial total pressure and temperature (AARPT12.5 and 

AARTT12.5). The bypass total pressure and total temperature ratios are the ratios of the averaged 

total values to the fan inlet reference values, 

RAARPT12 = AARPT12.5/PTo, 

and, 

RAARTT12 = AARTT12.5/TTo. 

The bypass duct adiabatic efficiency is then calculated from these ratios as, 

EADAAR12 = (RAARPT122/7-1)/(RAARTT12-1). 

For the core inlet total pressure, the pairs of total pressures were averaged at each of the five 

annular radii in the core inlet (ARPTCI (1 to 5)). Each of the five averaged total pressures was then 

used with the averaged core inlet static pressure to calculate a local Mach number. The total 

temperature recovery ratio for the core probes had not been calibrated. Instead, the recovery ratio 

was predicted based on the aspirated thermocouple geometry (ref. 11). The geometry (aspiration 

area) was then adjusted slightly to make the corrected temperature level consistent with the 

bypass duct temperatures. The indicated temperature of each thermocouple was then corrected by 

the temperature recovery ratio corresponding to the calculated local Mach number at each radius. 

The pairs of total temperatures were also averaged at each of the five annular radii in the core inlet 

(ARTTCI (1 to 5)). Then, the five annular averages of total pressure and temperature were 

averaged (based on equal areas) to calculate the annular averaged radial core inlet total pressure 

and total temperature (AARPTCI and AARTTCI). The core inlet total pressure and temperature 

ratios are the ratios of the core inlet averaged total values to the fan inlet reference values 

(AARPTCI/PTo and AARTTCI/TTo). 

The total fan pressure and temperature ratios (combining the bypass and core values) were 

calculated as the weight-flow averages of the bypass and core values. With the bypass and core 

flows, the weighted averages were, 

Fan Ratio = ((Bypass Duct Ratio * WBc)+(Core Inlet Ratio * WCc))/(WFc), 
NASA/TM—2006-213863 5

where WBc, WCc, and WFc are the bypass, core, and total fan weight flows, respectively, 

corrected to standard day conditions. 

The fan inlet total weight flow calculation used a Computational Fluid Dynamic (CFD) analysis 

of the inlet to predict the actual flow, including viscous effects. The freestream total pressure and 

local surface static pressure at the weight flow static pressure location (from the CFD solution) 

were used to calculate an “ideal” local Mach number and a correlating weight flow parameter 

(WFP). The WFP was chosen to be the Mach number part of the weight flow equation, 

WFcideal = Constant * PSTD/(TSTD)

0.5 * Area * (Mlocal/(1.0+0.2*Mlocal

2

)

3

), 

and, 

WFP = Mlocal/(1.0+0.2*Mlocal

2

)

3

. 

So, for totally uniform flow, the corrected flow would be a constant slope times the WFP value. 

Correlating the predicted actual flow (with boundary layer and static pressure gradient effects) with 

WFP made the correlation nearly a straight line, 

WFcCFD = fcorrelation(WFPCFD). 

The averaged cruciform rake measurements of freestream PTo and TTo and the averaged 

static pressure at the inlet weight flow static pressure location were used to calculate the actual 

corrected fan weight flow for each wind tunnel data point. The freestream total pressure and the 

local static pressure were used to calculate the “ideal” local Mach number and WFP. Then, the 

actual total fan weight flow corrected to standard day conditions (WFc) was calculated from the 

WFP and the CFD-predicted flow correlation. 

The core weight flow was calculated using a method similar to the fan inlet total weight flow. 

The flowfield was predicted along the core duct, back to the fully developed flow at the weight flow 

rake near the core nozzle. The total pressures from the CFD solution at the radii of the actual rake 

were averaged to get the averaged core flow total pressure. Similarly, the total temperatures from 

the CFD solution were also averaged. The total pressure and the averaged inner and outer duct 

static pressures were used to determine the “ideal” core flow rake Mach number and the core 

WFP, which were then used to correlate the predicted actual flow, corrected to the local core total 

pressure and total temperature as, 

WCclocal= WCCFD * (TTCW/TSTD)

0.5 / (PTCW/PSTD). 

Each wind tunnel data point used the core flow rake averaged total pressure, core inlet 

averaged total temperature, and averaged static pressure at the core weight flow location. The 

total pressure and the local static were used to calculate the “ideal” local Mach number and the 

WFP. The actual core weight flow corrected to local conditions (WCclocal) was calculated from the 

“ideal” local WFP and the predicted flow correlation as, 

WCclocal= fcorrelation(WFP). 

Then, the actual flow corrected to standard day was calculated using PTCW/PSTD and TTCI/TSTD

as, 

WCc = WCclocal * (PTCW/PSTD)/(TTCI/TSTD)

0.5. 

The bypass weight flow corrected to standard day conditions was calculated as the difference 

between the fan total corrected flow and the core corrected flow, 

WBc = WFc - WCc. 
NASA/TM—2006-213863 6

Fan Blade Configurations 

The fan blade design intent was to simulate a low tip speed, variable pitch (stagger) ADP-type 

fan. The takeoff tip speed and rpm, corrected to standard day conditions, were 840 ft/sec and 

8,750 rpm, respectively. The design is described in reference 12. Table 1 is a summary of fan 

blade design parameters and operating point parameters at the three noise measurement points: 

takeoff, approach, and cutback. The blade was designed as a graphite composite (GC) blade on a 

Titanium spar and a base button. The spar base was attached to the shank within the hub with a 

pin that allowed the blade to pivot in the flatwise direction (pin-root). At the base of the shank was a 

gear segment that allowed the fan blades to be locked and tested at different pitch angles. The 

cruise design angle was defined as 0 (zero) degrees. The takeoff angle was closed (farther away 

from axial) by 9 degrees from the cruise design angle, and identified as –9 degrees. A third, lower 

angle of –15 degrees was also tested. 

A solid Titanium (TI) blade set was designed and fabricated to be an aerodynamically similar 

replacement for the composite pin-root fan. A circular dovetail attachment allowed the same split 

hub to be used. The hub was designed with a split across the fan rotation axis at the fan stacking 

axis (MS 170.000). Gear segments were attached to the base of the dovetail to allow the same 

blade angle change capability as the pin-root, GC blades. For structural tuning, the inboard blade 

thickness was increased starting at about 60 percent span. The thickness matched the GC design 

at the root. The maximum increase in thickness/chord ratio of 0.015 occurred from 10 to 20 percent 

span. The total camber in the hub region was decreased by less than 4 percent. These changes 

were predicted to have negligible effect on the aerodynamic performance of the fan. 

Two different rubstrips were tested, installed in the fan casing over the fan blade tips. To 

minimize possible tip rubs and damage to the blades, the tip clearance was made to be 

0.055 inches at fan takeoff operating speed. That would represent a worst-case, worn engine 

clearance, and would be a worst-case for noise and performance. The baseline rubstrip was 

completely smooth. The second rubstrip had a Fan Casing Treatment (FCT) installed, designed to 

increase fan stall margin. Fan tip flow downstream of the fan was recirculated forward and injected 

upstream of the fan at the tip. It was expected that there would be some acoustic and performance 

penalty with the FCT rubstrip installed. The fan hardware combinations that were tested (blade 

material, blade angle, and rubstrip) are summarized in table 2. 

Fan Operating Line Results and Repeatability, TI Blade, 

–9 Degrees, FCT Rubstrip 

Two aerodynamic data runs with nearly identical fan/nacelle hardware were performed back-toback with the flight inlet and flight nozzle installed. The TI blades were installed at the takeoff angle 

(–9 deg) along with the FCT rubstrip. For the second run, only a single, small inlet boundary layer rake 

was removed (from the inlet flow static pressure location). For both runs, data were taken at 0.20, 0.15, 

and 0.10 freestream Mach number. The total fan performance results (combining the bypass and core 

measurements) are presented in figure 5, with the first run plotted with open symbols and the second 

with solid symbols. The fan total pressure and temperature ratio operating lines (versus fan corrected 

flow) are shown in figures 5(a) and (b). The two runs are nearly indistinguishable. The adiabatic 

efficiency, calculated from the pressure and temperature ratios, is shown in figure 5(c). The results 

show the efficiency repeats within less than 1.0 percent. The bypass duct rake performance results are 

presented in figure 6. Since the bypass flow is about 90 percent of the total flow, the bypass duct 

results look very much like the fan total results. The fan bypass pressure ratio and temperature ratio 

operating lines (versus fan corrected flow) are shown in figures 6(a) and (b). The two runs are nearly 

indistinguishable, and just slightly lower than the fan total results. The adiabatic efficiency, shown in 
NASA/TM—2006-213863 7

figure 6(c), is slightly higher than the fan total efficiency (fig. 5(c)). At flow rates below about 

75 lbm/sec, the efficiency is about 92.0 percent. At higher flow the efficiency drops, and at 84 lbm/sec 

the efficiency is about 84.0 percent. To complete the documentation of the fan operating line 

performance, the bypass ratio (WBc/WCc) is shown in figure 6d. The bypass ratio varies from about 

9.0 at low flow rates to 10.8 at high flow rates. The two runs are nearly indistinguishable. 

For selected, repeated operating conditions at Mo = 0.20 (shown in fig. 7(a)), details of the radial 

distribution of pressure ratio, temperature ratio, and adiabatic efficiency from the bypass duct rakes are 

shown in figures 7(b) through (i), respectively. The rake measurements at each radius in the core inlet 

and bypass duct were averaged (eliminating bad measurements), and the results plotted versus the 

radius. The pressure ratio and temperature ratio plots show only slight differences. These differences 

are magnified in the adiabatic efficiency plots (figs. 7(d) to (i)), however. Figures 7(e) to (i) show the 

efficiency more clearly by separating the five different fan speeds. The efficiency differences were 

about 1.0 percent, and up to 2.5 percent at 5,000 RPMc near the fan tip. (Note: The temperature rise 

at 5,000 RPMc is only about 16 °F, so 2.5 percent efficiency corresponds to only a 0.6 °F temperature 

difference). At higher fan speeds the efficiencies were almost identical. Though the core 

instrumentation was not as good or as reliable as the bypass rakes, for these runs even the core 

repeatability is quite good. As noted above for figure 5(c), the process of averaging these 10 radial 

measurements results in bypass duct fan efficiency averages that differ by less than 1.0 percent. The 

peak efficiency at mid-span radii ranges from about 98.5 percent at high fan speed (takeoff) to about 

97.5 at lower fan speed (approach). These results agree very well with the fan design predictions at 

takeoff, cutback and approach conditions. 

TI Versus GC Fan Operating Line Results at 

–9 Degree Angle, Smooth Rubstrip 

Fan operating line performance data were recorded for both the TI and GC blades, at the 

–9 degree blade angle with the smooth fan rubstrip installed. A comparison of the performance is 

shown in figure 8 (Note: there were no data recorded at Mach 0.15 for the GC blades). The open 

symbols are the TI results and the solid symbols are the GC results. In figures 8(a) and (b), the 

operating line pressure ratio and temperature ratio are quite similar. Noticeable at Mach 0.20, the TI 

blades operated at a slightly higher corrected flow, pressure ratio, and temperature ratio than the GC 

blades (comparing points where the same fan speed was run). The adiabatic efficiency (fig. 8(c)) 

shows little difference between the TI and GC blades at lower fan speeds and flow at Mach 0.20. The 

efficiency of the GC blades falls off faster at high fan flow, resulting in the TI blades being about 

3.0 percent higher at the highest flow (85 lbm/sec). The TI blades appear up to 2.0 percent lower in 

adiabatic efficiency at 0.10 Mach number, however. Though designed to be aerodynamically the 

same, the difference in stiffness and weight may cause differences in shape and performance as 

speed and loading changes. The bypass ratio (fig. 8(d)) shows little difference at 0.20 Mach number, 

but a much lower bypass ratio for the TI blades at 0.10 Mach number. This is unexpected, since the TI 

blades are thicker in the hub region and might be expected to perform worse there and result in lower 

core flow (therefore higher bypass ratio). The TI blades seem to run differently or be more sensitive 

when running along the lower-flow operating line at Mach 0.10. 

For selected, repeated operating fan speeds at Mo = 0.20 (fig. 9(a)), details of the radial 

distribution of pressure ratio, temperature ratio, and adiabatic efficiency from the bypass duct rakes are 

shown in figures 9(b) through (i), respectively. The TI blade pressure ratio was higher from the mid 

span to the tip. The TI blade temperature ratio was higher in the mid-span, but not at the very tip. The 

resulting efficiency was nearly identical in the hub to mid-span region, about 98.0 to 99.0 percent in the 

mid-span. The TI blade efficiency was about 1.0 percent better near the tip, for the higher fan speeds. 

For both blades, the highest fan speed had lower efficiency by 2 to 5 percent near the fan tip. 
NASA/TM—2006-213863 8

For the same fan speeds at Mo = 0.10 (fig. 10(a)), details of the radial distributions are shown in 

figures 10(b) through (i). An instability of the TI blades prevented testing at he higher fan speeds (at the 

lower fan flow rates of the Mo = 0.10 operating line). The average pressure and radial pressure 

distributions show the TI blades had a lower pressure rise than the GC from mid-span to the tip as fan 

speed was increased (TI blades were higher at Mo = 0.20). This significant reduction (with the core 

region showing no change) explains the unexpected reduction in bypass ratio for the TI blades at 

Mo = 0.10, noted above. The temperature at mid-span was lower with the TI blades as well. The TI 

blade efficiency no longer matched the GC blade. At 5,000 RPMc, the TI blade efficiency was higher at 

mid-span but was lower at the tip. At higher fan speeds, the TI efficiency dropped from mid-span to the 

tip (which did not happen at Mo = 0.20). This was quite a large difference with only 3 to 4 percent lower 

fan weight flow reduction on the Mo = 0.10 operating line (below the Mo = 0.20 operating line). Up to 

7,740 RPMc, there was very little change in efficiency near the blade tip. The GC blades had reduced 

efficiency at the highest fan speed by 2 to 3 percent (somewhat less than at Mo = 0.20). 

TI Versus GC Fan Operating Line Results at 

–15 Degree Angle, Smooth Rubstrip 

Fan operating line performance data were also recorded for both the TI and GC blades, at a 

–15 degree blade angle setting with the smooth rubstrip installed. A comparison of the performance is 

shown in figure 11 (Note: there were no data recorded at Mach 0.15 for the GC blades and no data at 

Mach 0.20 for the TI blades). The open symbols are the TI results and the solid symbols are the GC 

results. In figures 11(a) and (b), the operating line pressure ratio and temperature ratio are quite 

similar. Where the same fan speed was run, the TI blades operated at a higher corrected flow, 

pressure ratio, and temperature ratio. The adiabatic efficiency (fig. 11(c)) shows little difference 

between the TI and GC blades at moderate fan speeds and flow. At Mach 0.10 and low flow the GC 

performance was a percent or so higher, while the TI blades were about 2.0 percent higher in 

efficiency at the highest speed and flow (75 lbm/sec). However, at 0.20 Mach number the GC blades 

had a 4.0 percent or greater penalty in efficiency at all speeds (flows). At –15 degrees, the GC blades 

appear more sensitive to increasing speed and flow that occurred going to 0.20 Mach number. Both 

the TI and GC blades show a reduction in efficiency as speed and flow were increased, whereas at the 

–9 degree settings, the efficiency was nearly constant until the higher speeds (flows) were reached. 

The bypass ratio (fig. 11(d)) shows a higher bypass ratio for the TI blades at 0.10 Mach number. The 

loading distribution (pressure rise) from the fan hub to tip may explain this. 

For selected, repeated operating conditions at Mo = 0.10 (fig. 12(a)), details of the radial 

distribution of pressure ratio, temperature ratio, and adiabatic efficiency from the bypass duct rakes are 

shown in figures 12(b) through (i), respectively. The TI pressure ratio was higher, especially near the 

tip. The TI core pressure ratio was just slightly higher with the –15 degree angle, which may explain 

the higher bypass ratio with the TI blades. The TI temperature ratio was higher in the mid-span, but not 

as much at the very tip. The differences are larger than occurred at the –9 degree angle setting (fig. 9). 

The resulting efficiency was nearly identical in the hub to mid-span region, about 98.0 to 99.0 percent 

in the mid-span. The TI performance was about 1.0 percent better near the tip, for the higher fan 

speeds. For both blades, the higher fan speeds had reduced efficiency by 7 to 12 percent near the fan 

tip (from 73.0 to 66.0 percent and 89.5 to 77.5 percent). This penalty was larger than the 2 to 3 and 

2 to 5 percent penalty (at Mo = 0.10 and 0.20, respectively) at the –9 degree angle setting with 

increased speed. At Mo = 0.10 the penalty at 7,740 RPMc was about 2 to 4 percent with the 

–15 degree angle, compared to no change at –9 degrees. 
NASA/TM—2006-213863 9

–9 and –15 Degree Angle Fan Operating Line Results, 

TI Blade, Smooth Rubstrip 

The change in blade angle from –9 degrees (intended takeoff angle) to –15 degrees reduces the 

pressure ratio and weight flow obtained for each rotor speed. This is similar to changing the blade 

design point, except the blade twist is not reoptimized for new operating conditions. A comparison of 

the performance with the TI blades and smooth rubstrip is shown in figure 13 (there were no data 

recorded at Mach 0.20 at the –15 degree blade angle). The open symbols are the –9 degree results 

and the solid symbols are the –15 degree results. In figure 13(a), the pressure ratios follow nearly 

identical operating lines. But, in figure 13(b), the temperature ratio is higher for the –15 degree setting, 

at higher fan flows. The adiabatic efficiency (fig. 13(c)) shows the result of the higher temperature ratio. 

The efficiency at the –15 degree angle is much lower than at –9 degrees for the higher fan flows. At 

the highest speed and flow, the –15 degree performance was 5 to 10 percent lower. The efficiency is 

about the same for both angles at low speed and flow. The bypass ratio (fig. 13(d)) shows a higher 

bypass ratio at –15 degrees. This would be expected, since the –15 degree angle setting closes off the 

core. 

Additional figures 13(e) and (f) show the pressure ratio and corrected fan weight flow versus 

corrected fan speed. At 8,750 RPMc (Mach 0.15) the pressure ratios are 1.285 and 1.231, for –9 and 

–15 degree angles, respectively. The corresponding corrected flows are 79.2 and 72.5 lb/sec. The 

pressure rise (e.g., (pressure ratio-1.0), or 0.285 and 0.231) was reduced about 19 percent, and the 

flow was reduced by 8 percent, by decreasing the blade angle by 6 degrees. 

Another way to look at the adiabatic efficiency figure 13(c) is to identify a lower-loading condition to 

call its “takeoff” point. At a weight flow of 55 to 60 lbm/sec, the efficiency is about equal to the 

–9 degree angle efficiency at 80 lbm/sec (its takeoff flow rate). The corresponding fan speed is about 

6,750 RPMc, or a 23 percent reduction, and pressure ratio is 1.13 to 1.14, or a pressure rise about 1⁄2 

that for the –9 degree takeoff point (i.e., 0.14/0.28 = 0.5). The engine noise could be reduced 

substantially, but at the cost of doubling the number of engines or the engine flow area to get the same 

thrust. 

For selected, repeated operating conditions at Mo = 0.15 (fig. 14(a)), details of the radial 

distribution of pressure ratio, temperature ratio, and adiabatic efficiency from the bypass duct rakes are 

shown in figures 14(b) through (i), respectively. The blade angle reduction to –15 degrees reduced the 

pressure ratio near the fan hub somewhat. The pressure ratio reduction near the fan tip was much 

larger. Using the RPMc = 8,750 curves, there was about 8 percent reduction in pressure rise (i.e., 

pressure ratio minus one) near the hub and about 26 percent reduction near the tip. The reduction in 

temperature rise was similar. The adiabatic efficiency was somewhat higher at the fan hub with the 

–15 degree angle, and was more dependent on fan speed than with the –9 degree angle. Similarly, the 

efficiency is much lower at the fan tip and very dependent on fan speed, with about 9.5 percent 

efficiency loss from 5,000 to 8,750 RPMc, compared to about 1.0 percent loss with the –9 degree 

blade angle. 

–9 and –15 Degree Angle Fan Operating Line Results, 

GC Blade, Smooth Rubstrip 

The GC blades were also tested at –9 and –15 degree angles. A comparison of the –9 and 

–15 degree angle performance with GC blades is shown in figure 15 (there were no data recorded at 

Mach 0.15 for either blade angle). The open symbols are the –9 degree results and the solid symbols 

are the –15 degree results. Overall, the results are the same as with the TI blades shown in figure 13. 

In figure 15(a), the pressure ratios follow a nearly identical operating line, as occurred with the TI 
NASA/TM—2006-213863 10

blades. A slightly higher pressure ratio (at equal weight flows) can be seen for the –15 degree angle at 

the highest flows. In figure 15(b), the temperature ratio is definitely higher for the –15 degree setting, at 

higher fan flows. The adiabatic efficiency (fig. 15(c)) shows the result of the higher temperature ratio. At 

Mach 0.10 the efficiency is high at low weight flow but decreases rapidly at high flow, almost the same 

as the TI blades at Mach 0.10 and 0.15 (fig. 13(c)). At the highest speed and flow the –15 degree 

angle performance was 5 to 10 percent lower than with the –9 degree angle. At Mach 0.20, the 

efficiency is even low at the lower flows, not like the data at the other Mach numbers, where the 

efficiency remained near 94 percent. The bypass ratio (fig. 15(d)) shows a higher bypass ratio with the 

–15 degree angle. However, the bypass ratios are closer at low flows for the GC blades and farther 

apart at low flows for the TI blades. 

Additional figures 15(e) and (f) show the pressure ratio and corrected fan weight flow versus 

corrected fan speed. At 8,750 RPMc (Mach 0.10) the pressure ratios are 1.282 and 1.226, for –9 and 

–15 degree angles, respectively. The corresponding corrected flows are 78.2 and 71.1 lbm/sec. As 

was previously seen for the TI blades (at Mach 0.15), the pressure rise was reduced about 20 percent, 

and the flow was reduced by 9 percent, by decreasing the blade angle by 6 degrees. At Mach 0.20 the 

pressure rise was reduced about 22 percent and the flow was reduced by 10 percent with the 6 degree 

blade angle reduction. 

For selected, repeated operating conditions at Mo = 0.10 (fig. 16(a)), details of the radial 

distribution of pressure ratio, temperature ratio, and adiabatic efficiency from the bypass duct rakes are 

shown in figures 16(b) through (i), respectively. As with the TI blades at 0.15, the blade angle reduction 

to –15 degrees reduced the pressure ratio near the fan hub somewhat and much larger near the fan 

tip. Using the 8,750 RPMc curves, there was about 8 percent reduction in pressure rise (i.e., pressure 

ratio minus one) near the hub and about 25 percent reduction near the tip. The temperature ratio 

reduction was similar. As with the TI blades, the efficiency is somewhat higher at the fan hub with the 

–15 degree angle, and more dependent on fan speed than with the –9 degree angle. Similarly, the 

efficiency is much lower at the fan tip and very dependent on fan speed, with about 9.5 percent 

efficiency loss from 5,000 to 8,750 RPMc, compared to about 1.0 percent loss with the –9 degree 

blade angle. 

For similar repeated operating conditions at Mo = 0.20 (fig. 17(a)), details of the radial distribution 

of pressure ratio, temperature ratio, and adiabatic efficiency from the bypass duct rakes are shown in 

figures 17(b) through (i), respectively. The reductions with the –15 degree blade angle were similar, 

and just a little larger than at Mo = 0.10. Using the 8,750 RPMc curves, there was about 10 percent 

reduction in pressure rise near the hub and about 29 percent reduction near the tip. The overall 

efficiency was lower with the –15 degree angle at Mach 0.20, and the variation with fan speed was 

less than at Mach 0.10. Similarly, the efficiency was much lower at the fan tip but less dependent on 

fan speed than at Mach 0.10, with about 6.0 percent efficiency loss from 5,000 to 8,750 RPMc, 

compared to about 3.5 percent loss with the –9 degree blade angle. 

Smooth Versus FCT Rubstrip Operating Line Results, 

TI Blade, –9 Degree Angle 

Fan operating line performance data were recorded with the smooth rubstrip and with the FCT 

rubstrip installed, with the TI blades at the –9 degree blade angle. A comparison of the performance is 

shown in figure 18. The open symbols are with the smooth rubstrip and the solid symbols are with the 

FCT rubstrip. In figure 18(a), the operating line pressure ratios are similar, except the ratios with FCT 

rubstrip were noticeably lower at higher flow and pressure ratio (comparing points where the same fan 

speed was run, which are most of the nearby points). The corresponding temperature ratios (fig. 18(b)) 

appear similar to the pressure ratios. However, the temperature ratios for the smooth and FCT rubstrip 

were about the same at common fan speeds (or slightly lower with the FCT). The adiabatic efficiency 
NASA/TM—2006-213863 11

(fig. 18(c)) is correspondingly lower, by about 3.0 percent, with the FCT installed, even at the lower 

flow rates. The bypass ratio (fig. 18(d)) is also noticeably lower with the FCT rubstrip installed. The 

lower pressure ratio in the bypass duct, without a corresponding reduction in the core pressure ratio, 

would be expected to have this result. 

For selected, repeated operating conditions at Mo = 0.20 (fig. 19(a)), details of the radial 

distribution of pressure ratio, temperature ratio, and adiabatic efficiency from the bypass duct rakes are 

shown in figures 19(b) through (i), respectively. The pressure ratios with the smooth and FCT rubstrips 

were equal at the core and out to mid-way in the bypass duct. The FCT rubstrip caused a significantly 

reduced pressure ratio approaching the tip of the fan. The temperature ratios were equal at the core 

and out to mid-way in the bypass duct, also. The FCT rubstrip caused a reduced temperature ratio 

approaching the tip of the fan, except very close to the tip. There, re-circulating flow of the FCT 

apparently was reheated and showed up as an increase in temperature at the tip, even though 

pressure ratio was lower at the tip. The resulting efficiency was similar in the hub to mid-span region, 

about 97.0 to 99.0 percent in the mid-span. The FCT performance penalty was located almost 

completely in the top 20 percent of the fan annular area, with a loss of about 9 to 11 percent at the 

tip, and 6 to 8 percent at the next radius inboard. The FCT design intent was to have significantly less 

re-circulating flow than actually occurred. That would have resulted in much less performance penalty 

for the FCT. For the smooth rubstrip, the highest fan speed lost about 2.5 to 5.5 percent in efficiency at 

the last two radii near the fan tip, compared to the lowest speeds. For the FCT rubstrip, that loss was 

about 2.0 to 7.0 percent in efficiency. 

Fan Performance Map Results and Repeatability, GC Blade, 

–9 Degrees, Smooth Rubstrip 

The performance data presented in this report, so far, have been on operating lines with a flighttype inlet and fixed-area flight-type nozzle installed. One of the fan blades (GC blades at –9 degree 

angle, smooth rubstrip) was tested very early in the test schedule (Run 5) in a full fan mapping 

configuration and much later in an operating line configuration (Run 40) at the end of the performance 

phase of the tunnel entry. The mapping configuration was with a bellmouth inlet replacing the flight 

inlet. A Variable Fan Exit Nozzle (VFEN) replaced the fixed-area flight nozzle, with radial blades 

operating like a butterfly valve to change the exit area. For the mapping runs, the freestream Mach 

number was set nominally at 0.05 Mach number to improve the inlet temperature uniformity. Each fan 

speed was set and data recorded with the VFEN set at a large area. Initially there was no flow through 

the core (the lowest pressure ratio and highest flow points). Reducing the VFEN area back-pressured 

the fan, forcing flow into the core, instead of separating off the back side of the fan hub. Further VFEN 

area reduction allowed performance mapping from cruise conditions, across the takeoff operating line, 

and toward fan stall. 

For the later operating line run (being used for this comparison), data were taken at freestream 

Mach numbers of 0.10 and 0.20, with the flight inlet and nozzle installed. The bypass rake 

performance results are presented in figure 20, with the mapping data (Run 5) plotted with open 

symbols and the Mach 0.10 and 0.20 operating line data (Run 40) with solid symbols. The fan bypass 

pressure ratio map and temperature ratio map (versus fan corrected flow) are shown in figures 20(a) 

and (b). The operating line points match the mapping curves almost identically. The bypass adiabatic 

efficiency is shown in figure 20(c). The results show the efficiency repeats within better than 

1.0 percent. At flow rates below about 75 lbm/sec, the efficiency is about 94.0 percent. At higher flow 

the efficiency drops, and at 83 lbm/sec the efficiency is about 91.0 percent. To complete the 

documentation of the fan operating line performance, the bypass ratio (Wbypass/Wcore) is shown in 

figure 20(d). The mapping run bypass ratio varies from about 6 to 15 at low fan speed and 8 to18 at 

high fan speed. The operating line bypass ratio only varies from 9 at low speed to 12 at high speed. 
NASA/TM—2006-213863 12

The bypass ratio changes in different ways for mapping and operating line runs. Decreasing VFEN 

area in mapping runs increases back-pressure in the bypass duct only (not core), decreasing the 

bypass flow and the bypass ratio. Decreasing freestream Mach number in operating line runs changes 

the exit pressure of both the bypass and the core. For the operating line data, the result was increased 

bypass ratio with reduced flow (from lower freestream Mach number). 

TI Versus GC Fan Map Results at –9 Degree Angle, Smooth Rubstrip 

Fan map performance data were recorded for both the TI and GC blades, at the –9 degree blade 

angle with the smooth fan rubstrip installed. A comparison of the performance is shown in figure 21. 

The open symbols are the TI results and the solid symbols are the GC results. In figures 21(a) and (b), 

the mapped bypass pressure ratio and temperature ratio are almost the same for the two blade 

designs near the 0.10 Mach number operating line (shown), but everywhere else the TI and GC curves 

are different. At high flow, the TI blade pressure ratio and temperature ratio were higher than that of the 

GC blades. At lower flow, the TI blade pressure ratio and temperature ratio dipped down lower than 

the GC blades. At flows just above the Mach 0.10 operating line, the TI and GC performance curves 

cross and, where Mach 0.20 sea level operating line would be, the TI blades operated at a slightly 

higher corrected flow, pressure ratio, and temperature ratio. This is the same result seen in the 

operating line comparison (figs. 8(a) and (b)). As noted for those figures, though designed to be 

aerodynamically the same, the difference in stiffness and weight may cause differences in shape 

(twist) and performance as speed and loading changes for the TI and GC blades. At flows higher than 

the operating line, the reduced aerodynamic loading may allow the GC blades to untwist to a lower 

incidence, further reducing the GC pressure ratio relative to the TI blades. At flows just below the 

operating line, the increased aerodynamic loading may force the GC blades to twist to a higher 

incidence, increasing the GC pressure ratio. At much lower flows the two blades approach the same 

performance. 

The bypass adiabatic efficiency (fig. 21(c)) and bypass ratio (fig. 21(d)) show differences between 

the TI and GC blades, also. For clarity, the efficiency curves were re-plotted separately for each fan 

speed in figure 22. For a given fan speed, the bypass ratio for the TI blade is higher than for the GC 

blades at low weight flow, but crosses over and becomes lower than the GC blades at higher flow. A 

check of the bypass and core flow values showed that the core flow was lower for the TI blades at 

lower fan flow and higher at high fan flow, which caused the bypass ratio differences. 

As noted above, the efficiency for each TI and GC blade speed has been plotted on a separate 

plot (figs. 22(a) through (e)). At 5,000 RPMc, the few data points with lower efficiency 

(near 50 lbm/sec) are those where the core was not flowing. For those points, the efficiency is slightly 

better for the TI blades. The TI blades have nearly the same efficiency as the GC blades near where 

the core started flowing. At lower flow the TI blade performance fell off to about 1.0 percent below the 

GC blade. At 6,500 RPMc, the TI blades have nearly the same efficiency as the GC blades. Those 

points were near where the core started flowing (conditions without core flow were not set with the TI 

blades). Data were not taken to lower flow for the TI blades because high blade stresses were seen. At 

7,740 RPMc, the two points with the highest flow were higher in efficiency than the GC blades by 

about 1.0 percent. The points near the 0.10 sea level operating line have the same efficiency. The 

other points at lower flow were lower than the GC blades by about 2.0 percent. At 8,290 (8,300 for GC) 

and 8,750 RPMc, similar TI efficiency benefit can be seen at flow higher than the 0.10 operating line 

(up to 2.0 percent higher), and efficiency penalties at flow below the 0.10 operating line (up to 

2.0 percent lower). The operating line results (fig. 8(c)) gave the TI blades up to 2.0 percent lower 

efficiency at 0.10 Mach number, but show little difference above that, and about 3.0 percent higher 

efficiency than the GC blades at the highest speed (85 lbm/sec). 
NASA/TM—2006-213863 13

The peak efficiencies for the TI and GC blades were, respectively, 94.5 and 95.5 at 5,000 RPMc, 

94.0+ and 95.0+ at 6,500 RPMc, 94.0 and 95.0 at 7,740 RPMc, 94.0- and 95.0- at 8,290 RPMc, and, 

92.5 and 94.5- at 8,750 RPMc. 

TI Versus GC Fan Map Results at –9 Degree Angle, FCT Rubstrip 

Fan map performance data were recorded for both the TI and GC blades, at the –9 degree blade 

angle with the FCT fan rubstrip installed. A comparison of the performance is shown in figure 23. The 

open symbols are the TI results and the solid symbols are the GC results. In figures 23(a) and (b), the 

mapped fan pressure ratio and temperature ratio are only slightly different at flows below and including 

the Mach 0.10 operating line (shown). However, at high flow, the TI pressure ratio and temperature 

ratio were higher than the GC blades (similar to the smooth rubstrip results). The Mach 0.20 operating 

line would be at slightly higher flow than for the Mach 0.10 line, and the TI blades already show a 

slightly higher pressure ratio and temperature ratio than the GC blades. This is the same result seen at 

Mach 0.20 in the smooth rubstrip operating line comparison (fig. 8). However, in this case with the FCT 

installed, the pressure and temperature ratios at lower flows did not dip to the low levels seen with the 

smooth rubstrip (fig. 21). 

The adiabatic efficiency (fig. 23(c)) and bypass ratio (fig. 23(d)) show differences between the TI 

and GC blades. For clarity, the efficiency curves were re-plotted separately for each fan speed in 

figure 24. For a given fan speed, the bypass ratio for the TI blade is slightly higher than for the GC 

blades at low weight flow, but crosses over and becomes lower than the GC blades at high flow. This 

is similar to the results with the smooth rubstrip and core weight flow differences are also the reason 

for the bypass ratio differences. 

As noted above, the efficiency for each TI and GC blade speed has been plotted on a separate 

plot (figs. 24(a) through (c)). At 7,740 RPMc, the two points with the highest flow were higher than the 

GC blades by less than 1.0 percent (this is 0.5 percent less improvement than with the smooth 

rubstrip). At the 0.10 sea level operating line, both blades have the same efficiency. The other points at 

slightly lower flow were only lower than the GC blades by about 0.5 percent (this is about 1.5 percent 

less penalty than with the smooth rubstrip). At the lowest flow the TI and GC efficiencies are nearly 

equal again rather than the 2.0 percent reduction with the smooth rubstrip. At 8,290 and 8,750 RPMc, 

similar benefits can be seen at flow higher than the 0.10 operating line and 1.0 percent to no reduction 

at flow below the 0.10 operating line. 

The peak efficiencies for the TI and GC blades were, respectively, 91.5 and 92.0 at 7,740 RPMc, 

91.0+ and 91.0+ at 8,290 RPMc, and, 90.0+ and 91.0 at 8,750 RPMc. The FCT rubstrip minimized the 

efficiency difference between the blades, but overall both blades had a lower efficiency than either 

blade with the smooth rubstrip. The penalty with the FCT rubstrip was about 2.0 percent at 

7,740 RPMc and about 3.0 percent at the higher fan speeds. 

Smooth Versus FCT Rubstrip Fan Map Results, 

GC Blade, –9 Degree Angle 

Fan map performance data were presented in the previous two sections of this report comparing TI 

and GC blades, firs with the smooth rubstrip and then with the FCT rubstrip. The smooth and FCT 

rubstrips are compared in this section, at the –9 degree blade angle, with the GC blade installed. A 

comparison of the performance is shown in figure 25. The open symbols are the smooth rubstrip 

results and the solid symbols are the FCT results. In figures 25(a) and (b), the mapped fan pressure 

ratio and temperature ratio are different. The pressure ratio and temperature ratio were lower with the 

FCT rubstrip than that with the smooth rubstrip. The differences were somewhat greater at higher 
NASA/TM—2006-213863 14

flows. At lower flow than the 0.10 Mach number operating line (shown), the pressure ratio and 

temperature ratio level off as flow is reduced. 

The adiabatic efficiency (fig. 25(c)) and bypass ratio (fig. 25(d)) show differences between the 

smooth and FCT rubstrips. For clarity, the efficiency curves were replotted separately for each fan 

speed in figure 26. For each fan speed, the bypass ratios matched at lower weight flows, but then the 

bypass ratio with the FTC rubstrip appears higher at the higher flows. A check of the bypass and core 

flow values showed that the core flows were nearly the same for both rubstrips at low fan flow, but the 

core flow was lower for the FCT rubstrip at high fan flow, which caused the bypass ratio differences. 

As noted above, the efficiency for each rubstrip and blade speed has been plotted on a separate 

plot (figs. 26(a) through (c)). The FCT rubstrip performed lower than the smooth rubstrip at all 

conditions. At 7,740 RPMc, the FCT rubstrip efficiency was lower than with the smooth rubstrip by 

2.5 to 4.0 percent, with the penalty at the lowest fan flow rate only 1.0 percent. At 8,300 RPMc 

(8,290 for FCT), the penalty range was 2.0 to 4.0 percent, with the penalty at the lowest fan flow rate 

just over 1.0 percent. At 8,750 RPMc the range was 3.0 to 6.0 percent, with the penalty at the lowest 

fan flow rate again only about 1.0 percent. 

The GC blades were used for acoustic testing at Mach 0.10, with the smooth and the FCT rubstrip. 

The operating speeds for the acoustic points of interest (takeoff, cutback, and approach) were 8,750, 

7,525, and 5,425 RPMc, respectively. The Mach 0.10 operating line with the smooth rubstrip was 

documented directly and is shown in figure 8. The Mach 0.10 operating line conditions with the FCT 

rubstrip required interpolation from the fan map shown in figure 25. The cutback speed also required a 

slight extrapolation to the lower fan speed (7,740 RPMc was the lowest speed mapped). The approach 

speed required using the operating line with the TI blades and the FCT rubstrip (fig. 5) and correcting 

for the slight change (at that low speed condition) from TI to GC blades. Tables 3 and 4 give the fan 

weight flow (corrected to standard day), bypass duct pressure ratio, temperature ratio, and adiabatic 

efficiency, for the GC blades with the smooth and FCT rubstrips, respectively. 

Smooth Versus FCT Rubstrip Fan Map Results, TI Blade, –9 Degree Angle 

Fan map performance data were recorded for both the smooth and FCT rubstrips, at the –9 degree 

blade angle, with the TI blade installed. A comparison of the performance is shown in figure 27. The 

open symbols are the smooth rubstrip results and the solid symbols are the FCT results. In 

figures 27(a) and (b), the mapped fan pressure ratio and temperature ratio are different. The pressure 

ratio and temperature ratio were lower with the FCT rubstrip than that with the smooth rubstrip, but the 

differences are smaller than with the GC blades (fig. 25). The differences between rubstrips were 

somewhat greater at higher flows. Near the 0.10 Mach number operating line (shown), the pressure 

ratio and temperature ratio level off as flow is reduced (and actually dip lower with the smooth rubstrip). 

The adiabatic efficiency (fig. 27(c)) and bypass ratio (fig. 27(d)) show differences between the 

smooth and FCT rubstrips. For clarity, the efficiency curves were re-plotted separately for each fan 

speed in figure 28. For each fan speed, the bypass ratios nearly matched at the lower flows but the 

bypass ratio curves appear higher for the FTC rubstrip at higher flows, as seen with the GC blades. A 

check of the bypass and core flow values showed that the core flows were nearly the same for both 

rubstrips at low fan flow, but the core flow was lower for the FCT rubstrip at high fan flow, which 

caused the bypass ratio differences. 

As noted above, the efficiency for each rubstrip and blade speed has been plotted on a separate 

plot (figs. 28(a) through (e)). The FCT rubstrip performed lower than the smooth rubstrip at all 

conditions. At 5,000 RPMc, the FCT rubstrip efficiency was lower than with the smooth rubstrip by 

1.0 to 1.5 percent. At 6,250 RPMc, the penalty was about 1.5 to 2.0 percent. At 7,740 RPMc, the 

penalty ranged from 1.5 to 3.0 percent, with the higher penalty at the higher fan flow rates. At 8,300 

and 8,750 RPMc, the penalty range increased to 1.5 to 5.0 percent, with the higher penalty again at 

the higher fan flows. 
NASA/TM—2006-213863 15

Smooth Versus FCT Rubstrip Fan Map Results, TI Blade, 

–15 Degree Angle 

Fan map performance data were recorded for both the smooth and FCT rubstrips, at the –15 

degree blade angle, with the TI blade installed. A comparison of the performance is shown in figure 29. 

The open symbols are the smooth rubstrip results and the solid symbols are the FCT results. In figures 

29(a) and (b), the mapped fan pressure ratio and temperature ratio are different. The differences are 

larger than with the –9 degree angle setting (figs. 27(a) and (b)). The FCT pressure ratio and 

temperature ratio were lower than with the smooth rubstrip. The differences in pressure and 

temperature ratio were greater at mid to high fan flows. At flow well below the 0.10 Mach number 

operating line (shown), the pressure ratio leveled off as flow was reduced with the smooth rubstrip, but 

not with the FCT, which continued to increase in pressure ratio. 

The adiabatic efficiency (fig. 29(c)) and bypass ratio (fig. 29(d)) show differences between the TI 

and GC blades. For clarity, the efficiency curves were re-plotted separately for each fan speed in 

figure 30. For each fan speed, the bypass ratios nearly matched at the lower flows but the bypass ratio 

curves appear higher for the FTC rubstrip at higher flows, as seen with the –9 degree angle. The 

bypass ratio reduction with the FCT is slightly greater with the –15 degree blade angle than with the 

–9 degree angle. 

As noted above, the efficiency for each fan speed has been plotted on a separate plot 

(figs. 30(a) through (e)). With the –15 degree blade angle, the data show an increased performance 

penalty associated with using the FCT rubstrip. At 5,000 RPMc, the FCT rubstrip efficiency was lower 

than with the smooth rubstrip by 2.0 to 4.0 percent (higher than the 1.0 to 1.5 percent for the –9 degree 

angle). At 6,250 RPMc, the penalty was also 2.0 to 4.0 percent (higher than the 1.5 to 2.0 percent for –

9 degrees). At 7,740 RPMc, the penalty ranged from 3.0 to 5.0 percent (higher than the 1.5 to 

3.0 percent with –9 degrees), with the higher penalty at the higher fan flow rates. At 8,290 RPMc, the 

penalty range increased to 3.5 to 6.0 percent (higher than the 1.5 to 5.0 percent with –9 degrees). And, 

at 8,750 RPMc the penalty range was 3.5 to 7.0 percent (higher than the 1.5 to 5.0 percent with 

–9 degrees), with the higher penalty again at the higher fan flows. 

Smooth Versus FCT Rubstrip Fan Map Results, TI Blade, 0 Degree Angle 

Fan map performance data were recorded for both the smooth and FCT rubstrips, at the 0 degree 

(design) blade angle, with the TI blade installed. A comparison of the performance is shown in 

figure 31. The open symbols are the smooth rubstrip results and the solid symbols are the FCT results. 

In figures 31(a) and (b), the mapped fan pressure ratio and temperature ratio are different, but mostly 

at the higher fan weight flows. The FCT pressure ratio and, to a lesser degree, temperature ratio were 

lower than with the smooth rubstrip. Near the 0.10 Mach number operating line (shown), there was not 

much difference. 

The adiabatic efficiency (fig. 31(c)) and bypass ratio (fig. 31(d)) show differences between the 

smooth and FCT rubstrips. For clarity, the efficiency curves were re-plotted separately for each fan 

speed in figure 32. For each fan speed, the bypass ratios nearly matched at the lower flows but the 

bypass ratio curves appear higher for the FTC rubstrip at higher flows, as seen with the other blade 

angles. The bypass ratio reduction with the FCT is slightly greater with the 0 degree blade angle than 

with the –9 degree angle. Also, the overall level of the bypass ratio appears higher with the 0 degree 

blade angle. 

As noted above, the efficiency for each fan speed has been plotted on a separate plot (figs. 32(a) 

through (d)). The FCT performed somewhat lower than the smooth rubstrip at all conditions. At 

5,000 RPMc, the penalty was about 0.5 to 1.0 percent. At 6,500 RPMc, the penalty was about 
NASA/TM—2006-213863 16

1.0 percent. At 7,740 RPMc, the penalty was about 1.0 to 2.0 percent. And, at the cruise design speed 

of 8,400 RPMc, the penalty was 1.5 to 2.5 percent. These FCT penalties with the cruise design blade 

angle are much less than with the –9 and –15 degree angles. 

TI Versus GC Single-Radius Map Results at 

–9 Degree Angle, Smooth Rubstrip 

The comparison of TI and GC blade fan maps, at the –9 degree blade angle, with the smooth 

rubstrip (fig. 21) showed an unusual drop or leveling off of the TI blade pressure ratio as the flow was 

reduced below the 0.10 Mach number operating line (shown). A further comparison of the performance 

is shown here in figure 33, where the averaged pressure ratios at individual radii near the blade tip are 

plotted. The open symbols are the TI results and the solid symbols are the GC results. In figures 33(a) 

through (c) pressure ratio data from the 10th (outer most), 9th, and 8th radii are shown. At high fan 

speeds, the TI pressure ratio shows an extreme drop as flow was reduced. The GC blades show a 

similar but lesser drop, and it occurred at a lower flow. At the 7th radius (fig. 33(d)) the TI pressure ratio 

drop was reduced, and the GC had very little drop. The plots show that an extensive portion of the TI 

blade changed its performance as the flow was reduced below the 0.10 Mach number operating line. 

The GC blade performance change was less and occurred at lower flows (further from the 0.10 Mach 

number operating line). 

As with the averaged pressure ratios, the averaged temperature ratios at individual radii near the 

blade tip are plotted in figure 34. The open symbols are the TI results and the solid symbols are the 

GC results. The temperature ratio plots for the 10th (outer most) radius (fig. 34(a)) is very much like the 

pressure ratio plot of figure 33(a). The TI blades show a temperature drop or leveling-off as the flow 

was reduced below the 0.10 Mach number operating line. However, the temperature difference 

between TI and GC blades disappears at the lowest flows (figs. 34(b) through (d)), while the pressure 

still shows a penalty for the TI blades (figs. 33(b) through (d)). 

The effect on the adiabatic efficiency of the different pressure and temperature ratio characteristics 

are shown in figure 35, where the averaged adiabatic efficiencies at individual radii near the blade tip 

are plotted. At high fan flow the efficiency of the TI blades is significantly higher than the GC blades 

near the tip (figs. 35(a) through (c)). At the reduced flows where the pressure and temperature ratio 

drops were noted, the efficiency of the TI blades drops dramatically below the GC blades. At the 7th 

radius (fig. 35(d)) the performance differences are similar to but smaller than at the outer three radii. 

To clarify TI and GC efficiency differences, the efficiency at each TI blade speed has been plotted 

on a separate plot (figs. 36(a) through (e)) for the 10th (outer most) measurement. At 5,000 RPMc, the 

TI efficiency was about 4.0 percent higher than the GC blades at higher flow. At lower flow, near peak 

efficiency and the 0.10 Mach number operating condition; the TI blades had nearly the same efficiency 

as the GC blades. At lower flow the TI performance fell off to about 1.0 percent below the GC blade. At 

6,500 RPMc, the efficiency was about 4.0 percent higher for the TI blades at higher flow. Data were 

not taken to lower flow for the TI blades because of the high blade stresses. At 7,740 RPMc, the two 

points with the highest flow were higher than the GC blades by about 5.5 percent. The points near the 

peak efficiency and the 0.10 Mach number operating condition have the same efficiency. The other 

points at lower flow were lower in efficiency than the GC blades by about 3.0 percent. At 8,290 and 

8,750 RPMc, the TI efficiency is as much as 8.0 percent higher than the GC efficiency at the highest 

fan flow. The points near the peak efficiency and the 0.10 Mach number operating condition have the 

same efficiency. The other points at lower flow were lower in efficiency than the GC blades by about 

3.0 percent. 
NASA/TM—2006-213863 17

TI Versus GC Single-Radius Map Results at 

–9 Degree Angle, FCT Rubstrip

The comparison of TI and GC blade fan maps, at the –9 degree blade angle, with the FCT rubstrip 

(fig. 23) showed very little drop of the TI blade pressure ratio as the flow was reduced below the 

0.10 Mach number operating line (shown), much different than the drop with the smooth rubstrip. A 

further comparison of the performance is shown here in figure 37, where the averaged pressure ratios 

at individual radii near the blade tip are plotted. The open symbols are the TI results and the solid 

symbols are the GC results. In figures 37a, b, c, and d, pressure ratio data from the 10th (outer most), 

9th, 8th, and 7th radii are shown. At high fan speeds, the TI pressure ratio shows a leveling off as flow 

was reduced, and some drop at the 8th and 7th radii. The GC blades show a similar leveling, but it 

occurred at a lower flow. The leveling was less noticeable at 7,740 RPMc than at the higher two fan 

speeds. Once the GC blade levels off the pressure ratio matches the TI blades almost exactly. The 

plots show that an extensive portion of the both blades changed performance as the flow was reduced 

below the 0.10 Mach number operating line. It occurred for the GC blade a lower fan flow. 

As with the averaged pressure ratios, the averaged temperature ratios at individual radii near the 

blade tip are plotted in figure 38. The open symbols are the TI results and the solid symbols are the 

GC results. These temperature ratio plots are somewhat like the pressure ratio plots of figure 37, but 

the differences are smaller. The TI blades have a higher temperature ratio at high fan flow, and it dips 

below the GC blades as the flow was reduced, especially for the 8th and 7th radii. With the FCT 

rubstrip, both the pressure difference and temperature difference between TI and GC blades 

disappears at the lowest flows. 

The effect on the adiabatic efficiency of the different pressure and temperature ratio characteristics 

are shown in figure 39, where the averaged adiabatic efficiencies at individual radii near the blade tip 

are plotted. At high fan flows the efficiency of the TI blades is significantly higher than the GC blades 

near the tip (figs. 39(a) through (c)). At the lower flows where the pressure and temperature ratio drops 

were noted, the efficiency of the TI blades drops dramatically below the GC blades. At the 7th radius 

(fig. 39(d)) the performance differences are similar to but smaller than at the outer three radii. 

To clarify TI and GC efficiency differences, the efficiency at each TI blade speed has been plotted 

on a separate plot (figs. 40(a) through (c)) for the 10th (outer most) measurement. At 7,740 RPMc, at 

the highest flow the TI blade efficiency was higher than the GC blades by about 3.5 percent. At lower 

flow, near peak efficiency and the 0.10 Mach number operating condition, the TI blades had nearly the 

same efficiency as the GC blades. For the next lower point (66 lbm/sec) the TI blade efficiency was 

about 2.0 percent lower than the GC. Then the lowest flow point for the TI blade is almost exactly 

equal to the GC efficiency. At 8,290 RPMc, at the highest flow the TI blade efficiency was higher than 

the GC blades by about 5.5 percent. At lower flow, near peak efficiency and the 0.10 Mach number 

operating condition, the TI blades had nearly the same efficiency as the GC blades. For the next lower 

point (69 lbm/sec) the TI blade efficiency was about 3.0 percent lower than the GC. Then the lowest 

flow point for the TI blade is almost exactly equal to the GC efficiency. At 8,750 RPMc, at the highest 

flow the TI blade efficiency was higher than the GC blades by about 7.5 percent. At lower flow, near 

peak efficiency and the 0.10 Mach number operating condition, the TI blades had nearly the same 

efficiency as the GC blades. For the next lower point (73 lbm/sec) the TI blade efficiency was about 

2.5 percent lower than the GC. And, again, the lowest flow point for the TI blade is almost exactly 

equal to the GC efficiency. 

Smooth and FCT Single-Radius Map Results, TI Blades, –9 Degree Angle 

The comparison of smooth and FCT rubstrip fan maps, at the –9 degree blade angle, with the GC 

blades (fig. 25) and TI blades (fig. 27) showed reduced pressure and temperature ratios, and 2.5 to 
NASA/TM—2006-213863 18

6.0 percent and 1.0 to 5.0 percent lower efficiency, respectively, with the FCT installed. A further 

comparison of the performance is shown here in figure 41, where the averaged pressure ratios at 

individual radii near the blade tip are plotted. The open symbols are the smooth rubstrip results and the 

solid symbols are the FCT results. In figures 41(a) through (d), pressure ratio data from the 10th (outer 

most), 9th, 8th, and 7th radii are shown. At low fan speed the smooth and FCT rubstrip pressure ratios 

look similar, with the FCT somewhat lower than the smooth rubstrip. At high rotor speeds, the pressure 

ratio is significantly lower for the FCT. The smooth rubstrip results show the unusual drop in the 

pressure ratio noted in the discussion of figures 21 and 33. There, the pressure ratio drop occurred at 

lower flow rates for the GC blades than for the TI blades. Here, in figure 41, the effect of the FCT 

rubstrip on the pressure ratio was to reduce the drop dramatically, at the 10th (outer most) and 9th 

radii. The reduction was not as large at the 8th radius, and there was not much difference at the 7th 

radius. Just as with the performance penalty, the FCT tip treatment affected the unusual performance 

change in the tip region only. 

As with the averaged pressure ratios, the averaged temperature ratios at individual radii near the 

blade tip are plotted in figure 42. The open symbols are the smooth rubstrip results and the solid 

symbols are the FCT results. The temperature ratio plot for the 10th radius (fig. 42(a)) is different from 

the pressure ratio plot of figure 41(a), in that the temperature ratios are higher with the FCT rubstrip 

(the pressure ratio was lower). For the three lower radii (figs. 42(b) through (d)) the temperature ratios 

were lower with the FCT than with the smooth rubstrip, as were the pressure ratios (figs. 41(b) through 

(d)). The temperature ratios do not have the sudden drop with reduced fan flow seen for the pressure 

ratios, but they do level off with the smooth rubstrip installed. 

The effect on efficiency of the different pressure and temperature ratio characteristics shown 

above, are shown in figure 43, where the averaged adiabatic efficiencies at the 10th (outer most) 

measurement radius near the blade tip is plotted. To clarify smooth and FCT rubstrip efficiency 

differences, the efficiency of each fan speed has been plotted on a separate plot (figs. 43(a) through 

(e)). At 5,000 RPMc, the FCT rubstrip efficiency near the fan tip was lower by 5 to 8 percent. At 

6,250 RPMc, the FCT rubstrip efficiency near the fan tip was lower by 8 to 11 percent. At 7,740 an 

8,300 RPMc, the FCT rubstrip efficiency near the fan tip was lower by 9 to 12 percent. At 8,750 RPMc, 

the FCT rubstrip efficiency near the fan tip was lower by 10 to 13 percent. This penalty could be 

reduced if the recirculating flow of the FCT were reduced. The FCT flow rate appeared to be well 

above that required for stall management. 

TI Blade Instability Occurrences 

Initial tests with the TI blades installed were fan mapping runs which included strain gages 

installed on the fan blades. Occasional automatic “VFEN safe” actions occurred due to high fan 

blade stresses, where the VFEN area would be opened rapidly to a “safe” condition. This was not 

considered important because the fast acting VFEN easily was able to open and reduce the blade 

strain. However, when tests on the acoustic operating lines were started with the flight inlet and 

fixed area nozzle installed, there was no way to quickly move out of high stress conditions except a 

rig emergency shutdown. There were several high stress instabilities that occurred and began to 

limit operation to well below the intended takeoff speed of 8,750 RPMc. Between two back-to-back 

wind tunnel runs, a single inlet boundary layer rake was removed from STA 2.0 directly in front of 

the fan (removing the slight flow disturbance), and the operating line testing was attempted with a 

completely clean inlet. The fan instabilities with speed still occurred. At the time there were theories 

proposed, but no definitive cause of the instability was identified. 

The tip radii map performance results shown in this report (e.g., plots at the outer four duct rake 

radii) show that a dramatic loading change occurred as the flow was driven down in the stall 

direction by closing the VFEN (along constant RPMc lines). The blades may show an oscillation in 

performance between high and low loading conditions where this loading change occurs. The data 
NASA/TM—2006-213863 19

points where the rapid change in pressure ratio at the 10th (outer most) radius occurred were 

identified for both blade designs (TI and GC) and both rubstrips (smooth and FCT) from figures 33 

and 37. 

Figure 44 shows the low flow and high flow (i.e., low performance and high performance) 

transition points as solid and open circles, for both blade designs with the smooth rubstrip installed. 

The GC blades are shown in black and the TI blades in red. Fan operating lines for Mach 0.10, 

0.15, and 0.20 are also shown. The GC transition points are at lower flow and farther away from 

the operating lines than the TI blades. Near 70 lbm/sec flow and 1.20 pressure ratio, the lower 

Mach number operating lines are actually between the transition points for the TI blades (but not 

for the GC blades). There were two fan instability points identified during testing for the TI blades. 

One was at Mach 0.10 and the other at Mach 0.125. They are plotted as blue triangles and are 

located right between to two transition data points for the TI blades. 

Figure 45 shows the low flow and high flow transition points as solid and open circles, for both 

blade designs with the FCT rubstrip installed. As above, the GC blades are shown in black and the 

TI blades in red. Fan operating lines for Mach 0.10, 0.15, and 0.20 are also shown. The GC 

transition points are at lower flow and farther away from the operating lines than the TI blades. 

Again, near 70 lbm/sec flow and 1.20 pressure ratio, the lower Mach number operating lines are 

actually between the transition points for the TI blades (but not for the GC blades). There were five 

fan instability points identified during testing for the TI blades. Two were at Mach 0.10, two at 

Mach 0.15 and one at Mach 0.20. They are plotted as blue triangles, squares, and a diamond, 

respectively. The Mach 0.10 fan instability points are near high-flow edge of the transition data 

points for the TI blades. The Mach 0.15 instability points are between the transition data points, 

near the high-flow edge. And, the Mach 0.20 instability point is located near the high-flow edge of 

the transition points. The FCT shifted the transition points to somewhat lower fan flows, but since 

the operating lines were still between the transition points for the TI blades, the FCT did not 

prevent the occurrence of fan blade instabilities. 

Summary of Results 

A large scale model representative of a low-noise, high bypass ratio turbofan engine was 

tested for acoustics and performance in the NASA Glenn 9- by 15-Foot Low-Speed Wind Tunnel. 

The low tip speed ADP-type fan, nacelle, with an un-powered, flow-through core passage (with 

core inlet guide vanes) were simulated. The fan duct and core inlet were instrumented with total 

pressure and total temperature rakes. Fan performance comparisons were made for repeatability, 

blade material (Titanium and Graphite/Composite), blade angle (–9 degrees, takeoff; –15 degrees, 

lower loading; 0 degrees, design), and fan rubstrip (smooth and Fan Casing Treatment). 

Pairs of operating line runs and fan mapping runs were compared and showed that averaged 

pressure and temperature ratio (versus fan flow rate) repeated very well, and adiabatic efficiency, 

from those measurements, repeated within better than 1.0 percent. Similar results were seen in 

comparisons of the radial distribution of pressure and temperature ratio, and efficiency (still within 

about 1.0 percent repeatability). 

The TI and GC blade operating line performance was documented during fixed area nozzle 

runs (with the smooth rubstrip installed and a –9 degree blade angle). The TI blades had slightly 

higher pressure and temperature ratio than the GC blades, at the same fan speeds, and the 

efficiency was up to 3.0 percent higher at the higher speeds, but up to 2.0 percent lower at lower 

speeds. This may be the result of slightly different shapes between the TI and GC blades as speed 

and blade loading were increased. The comparisons of the radial distributions of pressure and 

temperature ratio showed the TI performance changes (above GC at Mach 0.20 and below GC at 

Mach 0.20) were mostly from blade mid span to the tip (though the temperature difference 

disappeared at the tip). The efficiency was nearly the same from the hub through the mid span, 
NASA/TM—2006-213863 20

and the TI blade was only 1.0 percent higher near the tip. So, there were moderate performance 

differences between the TI and GC blades along the Mach 0.10 and 0.20 operating lines. 

The TI and GC blade operating line performance was also documented at a –15 degree blade 

angle (with the smooth rubstrip installed). With the lower –15 degree blade angle setting the trends 

were similar to the –9 degree data. However, the GC blades showed a 3.0 to 4.0 percent penalty at 

0.20 Mach number that was not there at –9 degrees. Also, the efficiency fell increasingly for both 

blade designs as speed and flow were increased. With the –9 degree angle the efficiencies were 

nearly constant until the higher speeds and flows were reached, then the efficiency dropped. If the 

lower flow, high efficiency point were taken as the takeoff condition for a new engine (i.e., reduced 

fan speed and pressure ratio), the noise would be reduced, but the number of engines or the 

engine flow area would have to double to get the same thrust. 

With the –15 degree angle (both blades), the radial distributions of pressure and temperature 

ratio were similar to the –9 degree data, with the increase for the TI blades a bit higher from the 

hub to the tip for –15 degrees. The efficiency distributions showed a penalty further from the tip for 

both blades. Also, at higher fan speeds, there was 7 to 12 percent lower efficiency near the tip 

(compared to mid span), a lot larger than the 2 to 5 percent at the –9 degree settings. Similarly, the 

efficiency was much lower near the fan tip and very dependent on fan speed, with about 9.5 percent 

efficiency loss from 5,000 to 8,750 RPMc, compared to about 1.0 percent loss with the –9 degree 

blade angle, at Mach 0.10. However, the loss with increased fan speed at Mach 0.20 was only 

about 6.0 percent, compared to about 3.5 percent loss with the –9 degree blade angle. 

Differences between the smooth and FCT rubstrips were documented along fan operating lines 

with the TI blades at the –9 degree angle. The pressure ratios were lower with the FCT rubstrip (at 

equal fan speed) and the temperature ratios were only a little lower. The corresponding efficiency 

was about 3.0 percent lower at all flow rates. The radial distributions showed the FCT pressure 

loss occurred from mid span to the tip. There was a temperature loss starting at mid span, also, but 

recirculating flow caused an increased temperature at the tip, even though the pressure ratio was 

lower at the tip. The efficiency distribution showed the loss (of 6 to 11 percent) was almost 

completely in the top 20 percent of the fan annular area. 

The TI and GC blade fan mapping performance was documented during VFEN nozzle runs 

with both the smooth and FCT rubstrips installed (at –9 degree blade angle). Near the 0.10 Mach 

number operating line, the TI and GC blade pressure and temperature ratios and efficiencies were 

similar (but lower with FCT). At higher fan flow than the operating line, the TI blades had higher 

pressure and temperature ratios, and higher efficiencies than the GC blades. At lower flow than the 

operating line, the TI pressure and temperature ratios and efficiencies dropped lower than the GC. 

The difference was much less with the FCT rubstrip. The TI blade efficiency benefit at higher flow 

was 1.0 to 2.0 percent, and 2.0 percent penalty at lower flow. The difference in performance away 

from the Mach 0.10 operating line may be caused by differing blade twist changes. The TI blade 

shape would be dominated by centrifugal loads, basically changing with fan speed. The light GC 

blades may respond to air loads and change twist differently, thereby changing the airfoil angle and 

loading. 

The smooth and FCT rubstrip mapping performance was compared at the –9 degree blade 

angle, with both the GC blades, and then the TI blades. With both blades the pressure ratio was 

lower with the FCT rubstrip than with the smooth rubstrip, and the same was true of the 

temperature ratio except at lower flow rates. The difference was greater with the GC blades than 

with the TI blades. The efficiency with the FCT rubstrip was lower than with the smooth rubstrip at 

all conditions. With the GC blades the penalty ranged from 2.5 to 4.0 percent at 7,740 RPMc, 2.0 

to 4.0 percent at 8,300 RPMc, and 3.0 to 6.0 percent at 8,750 RPMc. The corresponding penalty 

with the TI blades ranged from 1.0 to 1.5 percent at 5,000 RPMc, 1.5 to 3.0 percent at 7,740 

RPMc, and 1.5 to 5.0 percent at 8,300 and 8,750 RPMc (somewhat less penalty than with the GC 

blades). The higher penalties occurred at the higher flow rates for each fan speed. The smooth and 

FCT rubstrip mapping performance was also compared at the –15 degree and zero degree blade 
NASA/TM—2006-213863 21

angles, with the TI blade installed. The performance differences were similar to the –9 degree 

angle data, but the differences were larger with the –15 degree and significantly smaller with the 

zero degree blade angle. 

An unusual drop or leveling off of the TI blade pressure ratio occurred with the smooth rubstrip 

as the flow was reduced, and was investigated by plotting results at individual radii near the blade 

tip. The TI pressure ratio showed an extreme drop as flow was reduced, especially at high fan 

speeds. The GC blades show a similar but lesser drop, and it occurred at a lower flow. The 

temperature ratio results were similar, but the difference between the TI and GC blades 

disappeared at lower flows. At high fan flow and speed, the efficiency of the TI blades was 

significantly higher than the GC blades near the tip (up to 8.0 percent). At the reduced flow where 

the pressure and temperature drop was noted, the efficiency of the TI blades dropped dramatically 

below the GC blades (about 3.0 percent). 

The performance results of the GC and TI blades at individual radii near the blade tips were 

also investigated with the FCT rubstrip installed. There was very little drop of the TI blade pressure 

ratio as the flow was reduced below the Mach 0.10 operating line (quite different from the drop-off 

with the smooth rubstrip). The temperature ratio changes were similar to the pressure ratio, but 

smaller. The efficiency differences between the TI and GC blades were similar to the results with 

the smooth rubstrip (up to 7.5 percent high at high flow and 2.5 percent low at low flow). 

A comparison of the performance results at individual radii with the smooth and FCT rubstrips 

was made with the TI blades, at –9 degree blade angle. As noted above, the FCT rubstrip 

smoothed out the large drop in pressure ratio as flow was reduced. But, that was accompanied by 

a lower pressure ratio and higher temperature ratio overall. That resulted in much lower efficiency. 

(5.0 to 13.0 percent for the outermost radius). This penalty could be reduced if the recirculating 

flow of the FCT were reduced. The FCT flow rate appeared to be well above that required for stall 

management. 

The unusual changes in loading that were identified near the blade tip were related to several 

high blades stress instabilities that occurred while testing the TI blades. The operating conditions 

for the instabilities were shown to occur between two distinct fan flow boundaries where dramatic 

changes in the blade tip performance occurred. The high and low boundaries for the TI blades 

occurred near the fixed nozzle operating lines, while the GC blade boundaries were at lower fan 

flows and away from the fixed nozzle operating lines. This seems to be why the GC blades did not 

have the instabilities that occurred with the TI blades. 

As expected, the effect of the FCT rubstrip was a performance penalty. The amount of 

recirculating flow at the fan tip was larger than required for the desired stall margin improvement, 

however. Reducing the FCT flow would also reduce the resulting performance penalty. There was 

a somewhat surprising performance difference between the TI and GC blades, mostly away from 

the fixed-nozzle operating lines. Though both blades showed some indication of a flow transition 

near the blade tips as fan flow was reduced, the transition occurred near the fixed flight-type nozzle 

takeoff operating lines with the TI blades and resulted in high blade stress instabilities. 

Though the pressure and temperature ratios were low for this fan design, the techniques used 

to improve thermocouple measurement accuracy gave repeated data runs with the adiabatic 

efficiency agreeing within one percent. The measured fan adiabatic efficiency at takeoff was 

93.7 percent which matched the design intent well. 

References 

1. Society of Automotive Engineers, “In-Flight Thrust Determination,” SAE AIR 1703, Nov. 1985, 

Reaffirmed Dec. 1992. 

2. MIDAP Study Group, “Guide to In-Flight Thrust Measurements of Turbojets and Fan Engines,” 

AGARDOGRAPH AG-237, 1979. 
NASA/TM—2006-213863 22

3. Robert J. Jeracki, “Model Engine Performance Measurement from Force Balance 

Instrumentation,” NASA/TM—1998-208486, 1998. 

4. James Dittmar and David Elliot, “Some Acoustic Results from the Pratt and Whitney Advanced 

Ducted Propulsor–Fan 1” NASA/TM—1999-209049, 1999. 

5. E. Brian Fite, “Report of Fan Performance from Duct Rake Instrumentation on a 1.294 Pressure 

Ratio, 725 ft/sec Tip Speed Turbofan Simulator Using Vaned Passage Casing Treatment,” 

NASA/TM—2006-214241, 2006. 

6. Richard Woodward, David Elliot, Christopher Hughes, and J.J. Berton, “Benefits of Swept and 

Leaned Stators for Fan Noise Reduction,” NASA/TM—1998-208661, AIAA–99–0479, 2001. 

7. Christopher Hughes, “Aerodynamic Performance of Scale-Model Turbofan Outlet Guide Vanes 

Designed for Low Noise,” NASA/TM—2001-211352, AIAA–2002–0374, January 2002. 

8. John Gazzaniga, “Performance of Advanced Fan Exit Guide Vane Concepts for High Speed 

Fans,” AIAA–2002–0377, January 2002. 

9. Ronald H. Soeder, “NASA Lewis 9- by 15-Foot Low Speed Wind Tunnel Users Manual,” NASA 

TM–106247, 1993. 

10. Society of Automotive Engineers, “Gas Turbine Engine Performance Station Identification 

and Nomenclature,” ARP 755A, 1974. 

11. AGAARD, “Recommended Practices for Measurement of Gas Path Pressures and 

Temperatures for performance Assessment of Aircraft Turbine Engines and Components,” 

AGARD-AR no. 245, 1990. 

12. David E Hobbs, Robert J. Neubert, et. al., “Low noise Research Fan Stage Design,” NASA 

CR-195382, 1995. 

TABLE 1.—FAN DESIGN PARAMETERS 

Pressure Ratio (Bypass, Stage) 

SLTO 1.284 

Cruise 1.294 

Approach 1.077 

Cutback 1.209 

 

Corrected Fan Speed, RPMc 

SLTO 8,750 (104.2%) 

Cruise 8,400 (100.0%) 

Approach 5,000 (59.5%) 

Cutback 7,740 (92.1%) 

 

Corrected Fan Tip Speed, UcTIP (ft/sec) 

SLTO 840 

Cruise 806 

Approach 480 

Cutback 743 

 

Corrected Specific Weight Flow, WFc/A (lbm/sec/ft2) 

SLTO 36.9 

Cruise 42.5 

Approach 22.7 

Cutback 33.3 

 

Bypass Ratio-Cruise 13.3 

Blade Number 0.18 

Vane Number 0.45 

Hub/tip Ratio 0.426 

Diameter-Leading Edge 22.0 
NASA/TM—2006-213863 23

TABLE 2.—FAN MODEL CONFIGURATIONS 

Blade Design Blade Angle Rubstrip 

GC –9 degrees Smooth 

GC –9 degrees FCT 

GC –15 degrees Smooth 

TI –9 degrees Smooth 

TI –9 degrees FCT 

TI –15 degrees Smooth 

TI –15 degrees FCT 

TI 0 degrees Smooth 

TI 0 degrees FCT 

TABLE 3.—ACOUSTIC OPERATING CONDITIONS-GC BLADE WITH SMOOTH RUBSTRIP 

Condition RPMc WFc (lbm/s) RAARPT12 RAARTT12 EADAR12 

Takeoff 8,750 78.11 1.2816 1.07842 0.9368 

Cutback 7,525 68.17 1.2052 1.05782 0.9475 

Approach 5,425 49.43 1.0994 1.02915 0.9412 

TABLE 4.—ACOUSTIC OPERATING CONDITIONS-GC BLADE WITH FCT RUBSTRIP 

Condition RPMc WFc (1bm/s) RAARPT12 RAARTT12 EADAR12 

Takeoff 8,750 76.53 1.2685 1.07789 0.9026 

Cutback 7,525 66.74 1.1953 1.05712 0.9152 

Approach 5,425 49.08 1.0972 1.02904 0.9249 
NASA/TM—2006-213863 24

Figure 1(a).—ADP type engine model in the NASA Glenn 9- by 15-Foot 

Low-Speed Wind Tunnel. 

ROTATING FORCE BALANCE

STATIC FORCE BALANCE

CORE EXIT

BYPASS EXIT

Figure 1(b).—The 22-inch diameter fan engine model cross section. 
Figure 2(a).—ADP model cross section showing key instrumentation locations. 

Core inlet (IGV leading edge

NASA/TM—2006-213863 25
Figure 2(b).—ADP fan model with the inlet and some fan blades removed. 

NASA/TM—2006-213863 26
NASA/TM—2006-213863 27

PSiw,O

Ro

Ri

Figure 3(a).—Inlet weight flow instrumentation. 

PTcw,1

PTcw,2

PScw,O

PScw, I

Ro

Ri

Figure 3(b).—Core weight flow instrumentation. 

Inlet Flow Instrumentation

Static Pressures

STA 164.440

Radius 10.518 in

Angle, deg

0

60

120

150

170

190

210

240

300

Aft looking forward 

Total Pressures

STA 195.850

Rakes PTcw,1 PTcw,2

Angle 50 140

R R

Ro 5.700 5.700

5.608 5.608

5.420 5.420

5.224 5.224

5.021 5.021

4.809 4.809

Ri 4.700 4.700

Static Pressures

STA 195.850

PScw,O PScw, I

R 5.700 4.700

Angle Angle

45 45

135 135

225 225 Aft looking forward 315 315
NASA/TM—2006-213863 28

TT12,1

TT12,2

TT12,3

TT12,4

PT12,1

PT12,2

PT12,3

PT12,4

PS12,O

PS12, I

Ro

Ri

Figure 4(a).—Fan bypass duct instrumentation. 

Total Temperatures

STA 175.000

Rakes TT12,1 TT12,2 TT12,3 TT12,4

Angle 46 132 228 314

RRRR

Ro 11.320 11.320 11.320 11.320

11.119 11.119 11.119 11.119

10.685 10.685 10.685 10.685

10.233 10.233 10.233 10.233

9.760 9.760 9.760 9.760

9.263 9.263 9.263 9.263

8.737 8.737 8.737 8.737

8.177 8.177 8.177 8.177

7.577 7.577 7.577 7.577

6.924 6.924 6.924 6.924

6.204 6.204 6.204 6.204

Ri 5.810 5.810 5.810 5.810

Total Pressures

STA 175.209 Offset 0.355

at Angle 36.0

Rakes PT12,1 PT12,2 PT12,3 PT12,4

Angle 46 132 228 314

RRRR

Ro 11.320 11.320 11.320 11.320

11.119 11.119 11.119 11.119

10.685 10.685 10.685 10.685

10.233 10.233 10.233 10.233

9.760 9.760 9.760 9.760

9.263 9.263 9.263 9.263

8.737 8.737 8.737 8.737

8.177 8.177 8.177 8.177

7.577 7.577 7.577 7.577

6.924 6.924 6.924 6.924

6.204 6.204 6.204 6.204

Ri 5.810 5.810 5.810 5.810

Static Pressures

PS12,O PS12, I

STA 175.000

Radius 11.320

Angle STA R Angle

30 174.300 5.791 0

60 174.300 5.791 270

120 176.000 5.810 0

150 176.000 5.810 270

170

190

210

240

300

330

Aft looking forward 
NASA/TM—2006-213863 29

TTci,1

TTci,2

PTci,1

PTci,2

PSci, I

Ro

Ri

Figure 4(b).—Core inlet instrumentation. 

Total Temperatures

STA 173.844-173.946

Rakes TTci,1 TTci,2

R Angle Angle

Ro 5.345

5.236 38.574 307.148

5.085 50.002 318.576

4.929 61.430 330.004

4.767 72.858 341.432

4.601 84.286 352.860

Ri 4.531

Total Pressures

STA 173.844-3.946

Rakes PTci,1 PTci,2

R Angle Angle

Ro 5.345

5.236 32.860 301.434

5.085 44.288 312.862

4.929 55.716 324.290

4.767 67.144 335.718

4.601 78.572 347.146

Ri 4.531

Static Pressures

STA 173.966

PSci,O PSci, I

R None 4.494

Angle

45

135

225

315

Aft looking forward 
NASA/TM—2006-213863 30

(a) Fan averaged total pressure ratio versus total fan corrected flow. 

(b) Fan averaged total temperature ratio versus total fan corrected flow. 

Figure 5.—Fan operating line results and repeatability 

with –9 deg blade angle, FCT rubstrip. 

Open Run 36 

Solid Run 37 

Open Run 36 

Solid Run 37 
NASA/TM—2006-213863 31

(c) Fan averaged adiabatic efficiency versus total fan corrected flow. 

Figure 5.—An operating line results and repeatability 

with –9 deg blade angle, FCT rubstrip. 

Open Run 36 

Solid Run 37 
NASA/TM—2006-213863 32

(a) Bypass total pressure ratio versus total fan corrected flow. 

(b) Bypass total temperature ratio versus total fan corrected flow. 

Figure 6.—Bypass operating line results and repeatability 

with –9 deg blade angle, FCT rubstrip. 

Open Run 36 

Solid Run 37 

Open Run 36 

Solid Run 37 
NASA/TM—2006-213863 33

(c) Bypass adiabatic efficiency versus total fan corrected flow. 

(d) Bypass ratio versus total fan flow. 

Figure 6.—Bypass operating line results and repeatability 

with –9 deg blade angle, FCT rubstrip. 

Open Run 36 

Solid Run 37 

Open Run 36 

Solid Run 37 
NASA/TM—2006-213863 34

(a) Bypass pressure ratio versus total fan corrected weight flow. 

(b) Fan total pressure ratio radial profiles. 

Figure 7.—Radial profiles on fan operating line showing results 

and repeatability with –9 deg blade angle, 

FCT rubstrip, Mo = 0.20. 
NASA/TM—2006-213863 35

(c) Fan total temperature ratio radial profiles. 

(d) Fan adiabatic efficiency radial profiles. 

Figure 7.—Radial profiles on fan operating line showing results 

and repeatability with –9 deg blade angle, 

FCT rubstrip, Mo = 0.20. 
NASA/TM—2006-213863 36

(e) Fan adiabatic efficiency radial profiles, 5,000 RPMc. 

(f) Fan adiabatic efficiency radial profiles, 6,250 RPMc. 

Figure 7.—Radial profiles on fan operating line showing results 

and repeatability with –9 deg blade angle, 

FCT rubstrip, Mo = 0.20. 
NASA/TM—2006-213863 37

(g) Fan adiabatic efficiency radial profiles, 7,740 RPMc. 

(h) Fan adiabatic efficiency radial profiles, 8,300 RPMc. 

Figure 7.—Radial profiles on fan operating line showing results and 

repeatability with –9 deg blade angle, FCT rubstrip, Mo = 0.20. 
NASA/TM—2006-213863 38

(i) Fan adiabatic efficiency radial profiles, 8,750 RPMc. 

Figure 7.—Radial profiles on fan operating line showing results and 

repeatability with –9 deg blade angle, FCT rubstrip, Mo = 0.20. 
NASA/TM—2006-213863 39

(a) Bypass total pressure ratio versus total fan corrected flow. 

(b) Bypass total temperature ratio versus total fan corrected flow. 

Figure 8.—TI and GC blade bypass operating line results 

with –9 deg blase angle, smooth rubstrip. 
NASA/TM—2006-213863 40

(c) Bypass adiabatic efficiency versus total fan corrected flow. 

(d) Bypass ratio versus total fan corrected flow. 

Figure 8.—TI and GC blade bypass operating line results 

with –9 deg blade angle, smooth rubstrip. 
NASA/TM—2006-213863 41

(a) Bypass total pressure ratio versus total fan corrected flow. 

(b) Fan total pressure ratio radial profiles. 

Figure 9.—TI and GC blade radial profiles on fan operating line, 

with –9 deg blade angle, smooth rubstrip, Mo = 0.20. 
NASA/TM—2006-213863 42

(c) Fan total temperature ratio radial profiles. 

(d) Fan adiabatic efficiency radial profiles. 

Figure 9.—TI and GC blade radial profiles on fan operating line, 

with –9 deg blade angle, smooth rubstrip, Mo = 0.20. 
NASA/TM—2006-213863 43

(e) Fan adiabatic efficiency radial profiles, 5,000 RPMc. 

(f) Fan adiabatic efficiency radial profiles, 6,250 and 6,500 RPMc. 

Figure 9.—TI and GC blade radial profiles on fan operating line, 

with –9 deg blade angle, smooth rubstrip, Mo = 0.20. 
NASA/TM—2006-213863 44

(g) Fan adiabatic efficiency radial profiles, 7,740 RPMc. 

(h) Fan adiabatic efficiency radial profiles, 8,300 RPMc. 

Figure 9.—TI and GC blade radial profiles on fan operating line, 

with –9 deg blade angle, smooth rubstrip, Mo = 0.20. 
NASA/TM—2006-213863 45

(i) Fan adiabatic efficiency radial profiles, 8,750 RPMc. 

Figure 9.—TI and GC blade radial profiles on fan operating line, 

with –9 deg blade angle, smooth rubstrip, Mo = 0.20. 
NASA/TM—2006-213863 46

(a) Bypass total pressure ratio versus total fan corrected flow. 

(b) Fan total pressure ratio radial profiles. 

Figure 10.—TI and GC blade radial profiles on fan operating line, 

with –9 deg blade angle, smooth rubstrip, Mo = 0.10. 
NASA/TM—2006-213863 47

(c) Fan total temperature ratio radial profiles. 

(d) Fan adiabatic efficiency radial profiles. 

Figure 10.—TI and GC blade radial profiles on fan operating line, 

with –9 deg blade angle, smooth rubstrip, Mo = 0.10. 
NASA/TM—2006-213863 48

(e) Fan adiabatic efficiency radial profiles, 5,000 RPMc. 

(f) Fan adiabatic efficiency radial profiles, 6,250 and 6,500 RPMc. 

Figure 10.—TI and GC blade radial profiles on fan operating line, 

with –9 deg blade angle, smooth rubstrip, Mo = 0.10. 
NASA/TM—2006-213863 49

(g) Fan adiabatic efficiency radial profiles, 7,740 RPMc. 

(h) Fan adiabatic efficiency radial profiles, 8,300 RPMc. 

Figure 10.—TI and GC blade radial profiles on fan operating line, 

with –9 deg blade angle, smooth rubstrip, Mo = 0.10. 
NASA/TM—2006-213863 50

(i) Fan adiabatic efficiency radial profiles, 8,750 RPMc. 

Figure 10.—TI and GC blade radial profiles on fan operating line, 

with –9 deg blade angle, smooth rubstrip, Mo = 0.10. 
NASA/TM—2006-213863 51

(a) Bypass total pressure ratio versus total fan corrected flow. 

(b) Bypass total temperature ratio versus total fan corrected flow. 

Figure 11.—TI and GC blade bypass operating line results 

with –15 deg blade angle, smooth rubstrip. 
NASA/TM—2006-213863 52

(c) Bypass adiabatic efficiency versus total fan corrected flow. 

(d) Bypass ratio versus total fan corrected flow. 

Figure 11.—TI and GC blade bypass operating line results 

with –15 deg blade angle, smooth rubstrip. 
NASA/TM—2006-213863 53

(a) Bypass total pressure ratio versus total fan corrected flow. 

(b) Fan total pressure ratio radial profiles. 

Figure 12.—TI and GC blade radial profiles on operating line 

with –15 deg blade angle, smooth rubstrip, Mo = 0.10. 
NASA/TM—2006-213863 54

(c) Fan total temperature ratio radial profiles. 

(d) Fan adiabatic efficiency radial profiles. 

Figure 12.—TI and GC blade radial profiles on operating line 

with –15 deg blade angle, smooth rubstrip, Mo = 0.10. 
NASA/TM—2006-213863 55

(e) Fan adiabatic efficiency radial profiles, 5,000 RPMc. 

(f) Fan adiabatic efficiency radial profiles, 6,250 and 6,500 RPMc. 

Figure 12.—TI and GC blade radial profiles on operating line 

with –15 deg blade angle, smooth rubstrip, Mo = 0.10. 
NASA/TM—2006-213863 56

(g) Fan adiabatic efficiency radial profiles, 7,740 RPMc. 

(h) Fan adiabatic efficiency radial profiles, 8,300 RPMc. 

Figure 12.—TI and GC blade radial profiles on operating line 

with –15 deg blade angle, smooth rubstrip, Mo = 0.10. 
NASA/TM—2006-213863 57

(i) Fan adiabatic efficiency radial profiles, 8,750 RPMc. 

Figure 12.—TI and GC blade radial profiles on operating line 

with –15 deg blade angle, smooth rubstrip, Mo = 0.10. 
NASA/TM—2006-213863 58

(a) Bypass total pressure ratio versus total fan corrected flow. 

(b) Bypass total temperature ratio versus total fan corrected flow. 

Figure 13.—The –9 and –15 deg blade angle bypass operating line 

results with TI blades, smooth rubstrip. 
NASA/TM—2006-213863 59

(c) Bypass adiabatic efficiency versus total fan corrected flow. 

(d) Bypass ratio versus total fan corrected flow. 

Figure 13.—The –9 and –15 deg blade angle bypass operating line 

results with TI blades, smooth rubstrip. 
NASA/TM—2006-213863 60

(e) Bypass total pressure ratio versus fan corrected speed. 

(f) Total fan corrected flow versus fan corrected speed. 

Figure 13.—The –9 and –15 deg blade angle bypass operating line 

results with TI blades, smooth rubstrip. 
NASA/TM—2006-213863 61

(a) Bypass total pressure ratio versus total fan corrected flow. 

(b) Fan total pressure ratio radial profiles. 

Figure 14.—The –9 and –15 deg blade angle radial profiles on operating line 

with TI blades, smooth rubstrip, Mo = 0.15. 
NASA/TM—2006-213863 62

(c) Fan total temperature ratio radial profiles. 

(d) Fan adiabatic efficiency radial profiles. 

Figure 14.—The –9 and –15 deg blade angle radial profiles on operating line 

with TI blades, smooth rubstrip, Mo = 0.15. 
NASA/TM—2006-213863 63

(e) Fan adiabatic efficiency radial profiles, 5,000 RPMc. 

(f) Fan adiabatic efficiency radial profiles, 6,250 RPMc. 

Figure 14.—The –9 and –15 deg blade angle radial profiles on operating line 

with TI blades, smooth rubstrip, Mo = 0.15. 
NASA/TM—2006-213863 64

(g) Fan adiabatic efficiency radial profiles, 7,740 RPMc. 

(h) Fan adiabatic efficiency radial profiles, 8,300 RPMc. 

Figure 14.—The –9 and –15 deg blade angle radial profiles on operating line 

with TI blades, smooth rubstrip, Mo = 0.15. 
NASA/TM—2006-213863 65

(i) Fan adiabatic efficiency radial profiles, 8,750 RPMc. 

Figure 14.—The –9 and –15 deg blade angle radial profiles on operating line 

with TI blades, smooth rubstrip, Mo = 0.15. 
NASA/TM—2006-213863 66

(a) Bypass total pressure ratio versus total fan corrected flow. 

(b) Bypass total temperature ratio versus total fan corrected flow. 

Figure 15.—The –9 and –15 deg blade angle bypass operating line 

results with GC blades, smooth rubstrip. 
NASA/TM—2006-213863 67

(c) Bypass adiabatic efficiency versus total fan corrected flow. 

(d) Bypass ratio versus total fan corrected flow. 

Figure 15.—The –9 and –15 deg blade angle bypass operating line results 

with GC blades, smooth rubstrip. 
NASA/TM—2006-213863 68

(e) Bypass total pressure ratio versus fan corrected speed. 

(f) Total fan corrected flow versus fan corrected speed. 

Figure 15.—The –9 and –15 deg blade angle bypass operating line 

results with GC blades, smooth rubstrip. 
NASA/TM—2006-213863 69

(a) Bypass total pressure ratio versus total fan corrected flow. 

(b) Fan total pressure ratio radial profiles. 

Figure 16.—The –9 and –15 deg blade angle radial profiles on operating line 

with GC blades, smooth rubstrip, Mo = 0.10. 
NASA/TM—2006-213863 70

(c) Fan total temperature ratio radial profiles. 

(d) Fan adiabatic efficiency radial profiles. 

Figure 16.—The –9 and –15 deg blade angle radial profiles on operating line 

with GC blades, smooth rubstrip, Mo = 0.10. 
NASA/TM—2006-213863 71

(e) Fan adiabatic efficiency radial profiles, 5,000 RPMc. 

(f) Fan adiabatic efficiency radial profiles, 6,500 RPMc. 

Figure 16.—The –9 and –15 deg blade angle radial profiles on operating line 

with GC blades, smooth rubstrip, Mo = 0.10. 
NASA/TM—2006-213863 72

(g) Fan adiabatic efficiency radial profiles, 7,740 RPMc. 

(h) Fan adiabatic efficiency radial profiles, 8,300 RPMc. 

Figure 16.—The –9 and –15 deg blade angle radial profiles on operating line 

with GC blades, smooth rubstrip, Mo = 0.10. 
NASA/TM—2006-213863 73

(i) Fan adiabatic efficiency radial profiles, 8,750 RPMc. 

Figure 16.—The –9 and –15 deg blade angle radial profiles on operating line 

with GC blades, smooth rubstrip, Mo = 0.10. 
NASA/TM—2006-213863 74

(a) Bypass total pressure ratio versus total fan corrected flow. 

(b) Fan total pressure ratio radial profiles. 

Figure 17.—The –9 and –15 deg blade angle radial profiles on operating line 

with GC blades, smooth rubstrip, Mo = 0.20. 
NASA/TM—2006-213863 75

(c) Fan total temperature ratio radial profiles. 

(d) Fan adiabatic efficiency radial profiles. 

Figure 17.—The –9 and –15 deg blade angle radial profiles on operating line 

with GC blades, smooth rubstrip, Mo = 0.20. 
NASA/TM—2006-213863 76

(e) Fan adiabatic efficiency radial profiles, 5,000 RPMc. 

(f) Fan adiabatic efficiency radial profiles, 6,500 RPMc. 

Figure 17.—The –9 and –15 deg blade angle radial profiles on operating line 

with GC blades, smooth rubstrip, Mo = 0.20. 
NASA/TM—2006-213863 77

(g) Fan adiabatic efficiency radial profiles, 7,740 RPMc. 

(h) Fan adiabatic efficiency radial profiles, 8,300 RPMc. 

Figure 17.—The –9 and –15 deg blade angle radial profiles on operating line 

with GC blades, smooth rubstrip, Mo = 0.20. 
NASA/TM—2006-213863 78

(i) Fan adiabatic efficiency radial profiles, 8,750 RPMc. 

Figure 17.—The –9 and –15 deg blade angle radial profiles on operating line 

with GC blades, smooth rubstrip, Mo = 0.20. 
NASA/TM—2006-213863 79

(a) Bypass total pressure ratio versus total fan corrected flow. 

(b) Fan total temperature ratio radial profiles. 

Figure 18.—Smooth and FCT rubstrip bypass operating line results, 

–9 deg blade angle, TI blade. 
NASA/TM—2006-213863 80

(c) Bypass adiabatic efficiency versus total fan corrected flow. 

(d) Bypass ratio versus total fan corrected flow. 

Figure 18.—Smooth and FCT rubstrip bypass operating line results, 

–9 deg blade angle, TI blade.
NASA/TM—2006-213863 81

(a) Bypass total pressure ratio versus total fan corrected flow. 

(b) Fan total pressure ratio radial profiles. 

Figure19.—Smooth and FCT rubstrip radial profiles on fan operating line, 

–9 deg blade angle, TI blade, Mo = 0.20.
NASA/TM—2006-213863 82

(c) Fan total temperature ratio radial profiles. 

(d) Fan adiabatic efficiency radial profiles. 

Figure19.—Smooth and FCT rubstrip radial profiles on fan operating line, 

–9 deg blade angle, TI blade, Mo = 0.20.
NASA/TM—2006-213863 83

(e) Fan adiabatic efficiency radial profiles, 5,000 RPMc. 

(f) Fan adiabatic efficiency radial profiles, 6,250 RPMc. 

Figure19.—Smooth and FCT rubstrip radial profiles on fan operating line, 

–9 deg blade angle, TI blade, Mo = 0.20.
NASA/TM—2006-213863 84

(g) Fan adiabatic efficiency radial profiles, 7,740 RPMc. 

(h) Fan adiabatic efficiency radial profiles, 8,300 RPMc. 

Figure19.—Smooth and FCT rubstrip radial profiles on fan operating line, 

–9 deg blade angle, TI blade, Mo = 0.20.
NASA/TM—2006-213863 85

(i) Fan adiabatic efficiency radial profiles, 8,750 RPMc. 

Figure19.—Smooth and FCT rubstrip radial profiles on fan operating line, 

–9 deg blade angle, TI blade, Mo = 0.20.
NASA/TM—2006-213863 86

(a) Bypass total pressure ratio versus total fan corrected flow. 

(b) Bypass total temperature ratio versus total fan corrected flow. 

Figure 20.—Bypass performance map results and repeatability 

with –9 deg blade angle, smooth rubstrip. 
NASA/TM—2006-213863 87

(c) Bypass adiabatic efficiency versus total fan corrected flow. 

(d) Bypass ratio versus total fan corrected flow. 

Figure 20.—Bypass performance map results and repeatability 

with –9 deg blade angle, smooth rubstrip. 
NASA/TM—2006-213863 88

(a) Bypass total pressure ratio versus total fan corrected flow. 

(b) Bypass total temperature ratio versus total fan corrected flow. 

Figure 21.—TI and GC blade bypass performance map 

results with –9 deg blade angle, smooth rubstrip. 
NASA/TM—2006-213863 89

(c) Bypass adiabatic efficiency versus total fan corrected flow. 

(d) Bypass ratio versus total fan corrected flow. 

Figure 21.—TI and GC blade bypass performance map results 

with –9 deg blade angle, smooth rubstrip. 
NASA/TM—2006-213863 90

(a) Bypass adiabatic efficiency versus total fan corrected flow. 

(b) Bypass adiabatic efficiency versus total fan corrected flow. 

Figure 22.—TI and GC blade single-speed bypass efficiency 

results with –9 deg blade angle, smooth rubstrip. 
NASA/TM—2006-213863 91

(c) Bypass adiabatic efficiency versus total fan corrected flow. 

(d) Bypass adiabatic efficiency versus total fan corrected flow. 

Figure 22.—TI and GC blade single-speed bypass efficiency results 

with –9 deg blade angle, smooth rubstrip. 
NASA/TM—2006-213863 92

(e) Bypass adiabatic efficiency versus total fan corrected flow. 

Figure 22.—TI and GC blade single-speed bypass efficiency results 

with –9 deg blade angle, smooth rubstrip. 
NASA/TM—2006-213863 93

(a) Bypass total pressure ratio versus total fan corrected flow. 

(b) Bypass total temperature ratio versus total fan corrected flow. 

Figure 23.—TI and GC blade bypass performance map results 

with –9 deg blade angle, FCT rubstrip. 
NASA/TM—2006-213863 94

(c) Bypass adiabatic efficiency versus total fan corrected flow. 

(d) Bypass ratio versus total fan corrected flow. 

Figure 23.—TI and GC blade bypass performance map results 

with –9 deg blade angle, FCT rubstrip. 
NASA/TM—2006-213863 95

(a) Bypass adiabatic efficiency versus total fan corrected flow. 

(b) Bypass adiabatic efficiency versus total fan corrected flow. 

Figure 24.—TI and GC blade single-speed bypass efficiency results 

with –9 deg blade angle, FCT rubstrip. 
NASA/TM—2006-213863 96

(c) Bypass adiabatic efficiency versus total fan corrected flow. 

Figure 24.—TI and GC blade single-speed bypass efficiency results 

with –9 deg blade angle, FCT rubstrip. 
NASA/TM—2006-213863 97

(a) Bypass total pressure ratio versus total fan corrected flow. 

(b) Bypass total temperature ratio versus total fan corrected flow. 

Figure 25.—Smooth and FCT bypass performance map 

results with –9 deg blade angle, GC blades. 
NASA/TM—2006-213863 98

(c) Bypass adiabatic efficiency versus total fan corrected flow. 

(d) Bypass ratio versus total fan corrected flow. 

Figure 25.—Smooth and FCT bypass performance map 

results with –9 deg blade angle, GC blades. 
NASA/TM—2006-213863 99

(a) Bypass adiabatic efficiency versus total fan corrected flow. 

(b) Bypass adiabatic efficiency versus total fan corrected flow. 

Figure 26.—Smooth and FCT single-speed bypass efficiency results 

with –9 deg blade angle, GC blades. 
NASA/TM—2006-213863 100

(c) Bypass adiabatic efficiency versus total fan corrected flow. 

Figure 26.—Smooth and FCT single-speed bypass efficiency results 

with –9 deg blade angle, GC blades. 
NASA/TM—2006-213863 101

(a) Bypass total pressure ratio versus total fan corrected flow. 

(b) Bypass total temperature ratio versus total fan corrected flow. 

Figure 27.—Smooth and FCT bypass performance map results 

with –9 deg blade angle, TI blades. 
NASA/TM—2006-213863 102

(c) Bypass adiabatic efficiency versus total fan corrected flow. 

(d) Bypass ratio versus total fan corrected flow. 

Figure 27.—Smooth and FCT bypass performance map results 

with –9 deg blade angle, TI blades. 
NASA/TM—2006-213863 103

(a) Bypass adiabatic efficiency versus total fan corrected flow. 

(b) Bypass adiabatic efficiency versus total fan corrected flow. 

Figure 28.—Smooth and FCT single-speed bypass efficiency results 

with –9 deg blade angle, TI blades. 
NASA/TM—2006-213863 104

(c) Bypass adiabatic efficiency versus total fan corrected flow. 

(d) Bypass adiabatic efficiency versus total fan corrected flow. 

Figure 28.—Smooth and FCT single-speed bypass efficiency results 

with –9 deg blade angle, TI blades. 
NASA/TM—2006-213863 105

(e) Bypass adiabatic efficiency versus total fan corrected flow. 

Figure 28.—Smooth and FCT single-speed bypass efficiency results 

with –9 deg blade angle, TI blades. 
NASA/TM—2006-213863 106

(a) Bypass total pressure ratio versus total fan corrected flow. 

(b) Bypass total temperature ratio versus total fan corrected flow. 

Figure 29.—Smooth and FCT bypass performance map TI blades. 
NASA/TM—2006-213863 107

(c) Bypass adiabatic efficiency versus total fan corrected flow. 

(d) Bypass ratio versus total fan corrected flow. 

Figure 29.—Smooth and FCT bypass performance map results with 

–15 deg blade angle, TI blades. 
NASA/TM—2006-213863 108

(a) Bypass adiabatic efficiency versus total fan corrected flow. 

(b) Bypass adiabatic efficiency versus total fan corrected flow. 

Figure 30.—Smooth and FCT single-speed bypass efficiency results 

with –15 deg blade angle, TI blades. 
NASA/TM—2006-213863 109

(c) Bypass adiabatic efficiency versus total fan corrected flow. 

(d) Bypass adiabatic efficiency versus total fan corrected flow. 

Figure 30.—Smooth and FCT single-speed bypass efficiency results 

with –15 deg blade angle, TI blades. 
NASA/TM—2006-213863 110

(e) Bypass adiabatic efficiency versus total fan corrected flow. 

Figure 30.—Smooth and FCT single-speed bypass efficiency results 

with –15 deg blade angle, TI blades. 
NASA/TM—2006-213863 111

(a) Bypass total pressure ratio versus total fan corrected flow. 

(b) Bypass total temperature ratio versus total fan corrected flow. 

Figure 31.—Smooth and FCT bypass performance map results 

with zero blade angle, TI blades. 
NASA/TM—2006-213863 112

(c) Bypass adiabatic efficiency versus total fan corrected flow. 

(d) Bypass ratio versus total fan corrected flow. 

Figure 31.—Smooth and FCT bypass performance map results 

with zero blade angle, TI blades. 
NASA/TM—2006-213863 113

(a) Bypass adiabatic efficiency versus total fan corrected flow, RPMc = 5,000. 

(b) Bypass adiabatic efficiency versus total fan corrected flow, RPMc = 6,500. 

Figure 32.—Smooth and FCT single-speed bypass efficiency 

results with 0 degree blade angle, TI blades. 
NASA/TM—2006-213863 114

(c) Bypass adiabatic efficiency versus total fan corrected flow, RPMc = 7,740. 

(d) Bypass adiabatic efficiency versus total fan corrected flow, RPMc = 8,400. 

Figure 32.—Smooth and FCT single-speed bypass efficiency results 

with 0 degree blade angle, TI blades. 
NASA/TM—2006-213863 115

(a) Bypass 10th radius total pressure ratio versus total fan corrected flow. 

(b) Bypass 9th radius total pressure ratio versus total fan corrected flow. 

Figure 33.—TI and GC blade single-radius pressure ratio map results 

with –9 deg blade angle, smooth rubstrip. 
NASA/TM—2006-213863 116

(c) Bypass 8th radius total pressure ratio versus total fan corrected flow. 

(d) Bypass 7th radius total pressure ratio versus total fan corrected flow. 

Figure 33.—TI and GC blade single-radius pressure ratio map results 

with –9 deg blade angle, smooth rubstrip. 
NASA/TM—2006-213863 117

(a) Bypass 10th radius total temperature ratio versus total fan corrected flow. 

(b) Bypass 9th radius total temperature ratio versus total fan corrected flow. 

Figure 34.—TI and GC blade single-radius temperature ratio map results 

with –9 deg blade angle, smooth rubstrip. 
NASA/TM—2006-213863 118

(c) Bypass 8th radius total temperature ratio versus total fan corrected flow. 

(d) Bypass 7th radius total temperature ratio versus total fan corrected flow. 

Figure 34.—TI and GC blade single-radius temperature ratio map results 

with –9 deg blade angle, smooth rubstrip. 
NASA/TM—2006-213863 119

(a) Bypass 10th radius adiabatic efficiency versus total fan corrected flow. 

(b) Bypass 9th radius adiabatic efficiency versus total fan corrected flow. 

Figure 35.—TI and GC blade single-radius adiabatic efficiency map results 

with –9 deg blade angle, smooth rubstrip. 
NASA/TM—2006-213863 120

(c) Bypass 8th radius adiabatic efficiency versus total fan corrected flow. 

(d) Bypass 7th radius adiabatic efficiency versus total fan corrected flow. 

Figure 35.—TI and GC blade single-radius adiabatic efficiency map results 

with –9 deg blade angle, smooth rubstrip. 
NASA/TM—2006-213863 121

(a) Bypass 10th radius adiabatic efficiency versus total fan 

corrected flow, RPMc = 5,000. 

(b) Bypass 10th radius adiabatic efficiency versus total fan 

corrected flow, RPMc = 6,500. 

Figure 36.—TI and GC blade single-speed, single-radius, efficiency results 

with –9 deg blade angle, smooth rubstrip. 
NASA/TM—2006-213863 122

(c) Bypass 10th radius adiabatic efficiency versus total fan 

corrected flow, RPMc = 7,740. 

(d) Bypass 10th radius adiabatic efficiency versus total fan 

corrected flow, RPMc = 8,300. 

Figure 36.—TI and GC blade single-speed, single-radius, efficiency 

results with –9 deg blade angle, smooth rubstrip. 
NASA/TM—2006-213863 123

(e) Bypass 10th radius adiabatic efficiency versus 

total fan corrected flow, RPMc = 8,750. 

Figure 36.—TI and GC blade single-speed, single-radius, efficiency 

results with –9 deg blade angle, smooth rubstrip. 
NASA/TM—2006-213863 124

(a) Bypass 10th radius total pressure ratio versus total fan corrected flow. 

(b) Bypass 9th radius total pressure ratio versus total fan corrected flow. 

Figure 37.—TI and GC blade single-radius pressure ratio map 

results with –9 deg blade angle, FCT rubstrip. 
NASA/TM—2006-213863 125

(c) Bypass 8th radius total pressure ratio versus total fan corrected flow. 

(d) Bypass 7th radius total pressure ratio versus total fan corrected flow. 

Figure 37.—TI and GC blade single-radius pressure ratio map results 

with –9 deg blade angle, FCT rubstrip. 
NASA/TM—2006-213863 126

(a) Bypass 10th radius total temperature ratio versus total fan corrected flow. 

(b) Bypass 9th radius total temperature ratio versus total fan corrected flow. 

Figure 38.—TI and GC blade single-radius temperature ratio map results 

with –9 deg blade angle, FCT rubstrip. 
NASA/TM—2006-213863 127

(c) Bypass 8th radius total temperature ratio versus total fan corrected flow. 

(d) Bypass 7th radius total temperature ratio versus total fan corrected flow. 

Figure 38.—TI and GC blade single-radius temperature ratio map 

results with –9 deg blade angle, FCT rubstrip. 
NASA/TM—2006-213863 128

(a) Bypass 10th radius adiabatic efficiency versus total fan corrected flow. 

(b) Bypass 9th radius adiabatic efficiency versus total fan corrected flow. 

Figure 39.—TI and GC blade single-radius adiabatic efficiency map 

results with –9 deg blade angle, FCT rubstrip. 
NASA/TM—2006-213863 129

(c) Bypass 8th radius adiabatic efficiency versus total fan corrected flow. 

(d) Bypass 7th radius adiabatic efficiency versus total fan corrected flow. 

Figure 39.—TI and GC blade single-radius adiabatic efficiency map 

results with –9 deg blade angle, FCT rubstrip. 
NASA/TM—2006-213863 130

(a) Bypass 10th radius adiabatic efficiency versus total fan 

corrected flow, RPMc = 7,740. 

(b) Bypass 10th radius adiabatic efficiency versus total fan 

corrected flow, RPMc = 8,300. 

Figure 40.—TI and GC blade single-speed, single-radius, efficiency 

results with –9 deg blade angle, FCT rubstrip. 
NASA/TM—2006-213863 131

(c) Bypass 10th radius adiabatic efficiency versus total fan 

corrected flow, RPMc = 8,750. 

Figure 40.—TI and GC blade single-speed, single-radius, efficiency 

results with –9 deg blade angle, FCT rubstrip. 
NASA/TM—2006-213863 132

(a) Bypass 10th radius total pressure ratio versus total fan corrected flow. 

(b) Bypass 9th radius total pressure ratio versus total fan corrected flow. 

Figure 41.—Smooth and FCT rubstrip single-radius pressure ratio map 

results with –9 deg blade angle, TI blades.
NASA/TM—2006-213863 133

(c) Bypass 8th radius total pressure ratio versus total fan corrected flow. 

(d) Bypass 7th radius total pressure ratio versus total fan corrected flow. 

Figure 41.—Smooth and FCT rubstrip single-radius pressure ratio map 

results with –9 deg blade angle, TI blades.
NASA/TM—2006-213863 134

(a) Bypass 10th radius total temperature ratio versus total fan corrected flow. 

(b) Bypass 9th radius total temperature ratio versus total fan corrected flow. 

Figure 42.—Smooth and FCT rubstrip single-radius temperature ratio map 

results with –9 deg blade angle, TI blades.
NASA/TM—2006-213863 135

(c) Bypass 8th radius total temperature ratio versus total fan corrected flow. 

(d) Bypass 7th radius total temperature ratio versus total fan corrected flow. 

Figure 42.—Smooth and FCT rubstrip single-radius temperature ratio map results 

with –9 deg blade angle, TI blades.
NASA/TM—2006-213863 136

(a) Bypass 10th radius adiabatic efficiency versus total fan 

corrected flow, RPMc = 5,000. 

(b) Bypass 10th radius adiabatic efficiency versus total fan 

corrected flow, RPMc = 6,250. 

Figure 43.—Smooth and FCT rubstrip single-speed, single-radius 

efficiency results with –9 deg blade angle, TI blades.
NASA/TM—2006-213863 137

(c) Bypass 10th radius adiabatic efficiency versus total fan 

corrected flow, RPMc = 7,740. 

(d) Bypass 10th radius adiabatic efficiency versus total fan 

corrected flow, RPMc = 8,300. 

Figure 43.—Smooth and FCT rubstrip single-speed, single-radius 

efficiency results with –9 deg blade angle, TI blades.
NASA/TM—2006-213863 138

(e) Bypass 10th radius adiabatic efficiency versus total fan 

corrected flow, RPMc = 8,750. 

Figure 43.—Smooth and FCT rubstrip single-speed, single-radius 

efficiency results with –9 deg blade angle, TI blades. 
Fan 1 Instability Boundaries

1.351.301.251.201.151.101.051.00

30.0 35.0 40.0 45.0 50.0 55.0 60.0 65.0 70.0 75.0 80.0 85.0 90.0

Corrected Fan Flow (WFc), lb/sec

Bypass Pressure Ratio, AARPT12/PTo

Fan OL, Mo= 0.100Fan OL, Mo= 0.125Fan OL, Mo= 0.150Fan OL, Mo= 0.200GC Low Trans PointGC High Trans PointTI Low Trans PointTI High Trans PointTI Instability, Mo= 0.100TI Instability, Mo= 0.125

TI Run 38 / Smooth

11/18/2004

Figure 44.—TI and GC blade operating lines, transition points, and instability points with smooth rubstrip.

NASA/TM—2006-213863 139
Fan 1 Instability Boundaries

1.351.301.251.201.151.101.051.00

30.0 35.0 40.0 45.0 50.0 55.0 60.0 65.0 70.0 75.0 80.0 85.0 90.0

Corrected Fan Flow (WFc), lb/sec

Bypass Pressure Ratio, AARPT12/PTo

Fan OL, Mo= 0.100Fan OL, Mo= 0.150Fan OL, Mo= 0.200GC Low Trans PointGC High Trans PointTI Low Trans PointTI High Trans PointTI Instability, Mo= 0.100TI Instability, Mo= 0.100rptTI Instability, Mo= 0.150TI Instability, Mo= 0.150rptTI Instability, Mo= 0.200

TI Runs 36&37 / FCT

11/18/2004

Figure 45.—TI and GC blade operating lines, transition points, and instability points with FCT rubstrip. 

NASA/TM—2006-213863 140
This publication is available from the NASA Center for AeroSpace Information, 301–621–0390.

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February 2006

NASA TM—2006–213863

E–15233

WBS–22–781–30–50

147

Comprehensive Report of Fan Performance From Duct Rake Instrumentation on

1.294 Pressure Ratio, 806 ft/sec Tip Speed Turbofan Simulator Models

Robert J. Jeracki

Wind tunnel; Powered model; Scale model; Turbofan; Ducted fan; Turbomachinery; Fan

noise; Low noise; Low tip speed

Unclassified -Unlimited

Subject Category: 02

Responsible person, Robert J. Jeracki, organization code RTA, 216–433–3917.

A large scale model representative of an advanced ducted propulsor-type, low-noise, very high bypass ratio turbofan engine

was tested for acoustics, aerodynamic performance, and off-design operability in the NASA Glenn 9- by 15-Foot Low-Speed

Wind Tunnel. The test was part of NASA’s Advanced Subsonic Technology Noise Reduction Program. The low tip speed

fan, nacelle, and un-powered core passage were simulated. As might be expected, the effect of stall management casing treatment was a performance penalty. Reducing the recirculating flow at the fan tip reduced the penalty while still providing

sufficient stall margin. Two fans were tested with the same aerodynamic design; one with graphite composite material, and

the other with solid titanium. There were surprising performance differences between the two fans, though both blades

showed some indication of transitional flow near the tips. Though the pressure and temperature ratios were low for this fan

design, the techniques used to improve thermocouple measurement accuracy gave repeatable data with adiabatic efficiencies

agreeing within 1 percent. The measured fan adiabatic efficiency at simulated takeoff conditions was 93.7 percent and

matched the design intent.