Daniel L. Tweedt

AP Solutions, Inc., Cleveland, Ohio

Preliminary Aerodynamic Investigation of

Fan Rotor Blade Morphing

NASA/CR—2012-217815

December 2012
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Daniel L. Tweedt

AP Solutions, Inc., Cleveland, Ohio

Preliminary Aerodynamic Investigation of

Fan Rotor Blade Morphing

NASA/CR—2012-217815

December 2012

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared under Contract NNC06BA07B, Task order NNC10E541
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Level of Review: This material has been technically reviewed by NASA technical management OR expert reviewer(s).
Abstract

Various new technologies currently under development may enable controlled blade shape 

variability, or so-called blade morphing, to be practically employed in aircraft engine fans and 

compressors in the foreseeable future. The current study is a relatively brief, preliminary 

computational fluid dynamics investigation aimed at partially demonstrating and quantifying the 

aerodynamic potential of fan rotor blade morphing. The investigation is intended to provide 

information useful for near-term planning, as well as aerodynamic solution data sets that can be 

subsequently analyzed using advanced acoustic diagnostic tools, for the purpose of making fan 

noise comparisons.

Two existing fan system models serve as baselines for the investigation: the Advanced 

Ducted Propulsor fan with a design tip speed of 806 ft/sec and a pressure ratio of 1.294, and the 

Source Diagnostic Test fan with a design tip speed of 1215 ft/sec and a pressure ratio of 1.470. 

Both are 22-inch sub-scale, low-noise research fan/nacelle models that have undergone extensive 

experimental testing in the 9- by 15-foot Low Speed Wind Tunnel at the NASA Glenn Research 

Center.

The study, restricted to fan rotor blade morphing only, involves a fairly simple blade 

morphing technique. Specifically, spanwise-linear variations in rotor blade-section setting angle 

are applied to alter the blade shape; that is, the blade is linearly retwisted from hub to tip. 

Aerodynamic performance comparisons are made between morphed-blade and corresponding 

baseline configurations on the basis of equal fan system thrust, where rotor rotational speed for the 

morphed-blade fan is varied to change the thrust level for that configuration.

The results of the investigation confirm that rotor blade morphing could be a useful 

technology, with the potential to enable significant improvements in fan aerodynamic performance. 

Even though the study is very limited in scope and confined to simple geometric perturbations of 

two existing fan systems, the aerodynamic effectiveness of blade morphing is demonstrated by the 

configurations analyzed. In particular, for the Advanced Ducted Propulsor fan it is demonstrated 

that the performance levels of the original variable-pitch baseline design can be achieved using 

blade morphing instead of variable pitch, and for the Source Diagnostic Test fan the performance at 

important off-design operating points is substantially increased with blade morphing.

NASA/CR—2012-217815 1

Preliminary Aerodynamic Investigation of 

Fan Rotor Blade Morphing 

Daniel L. Tweedt 

AP Solutions, Inc. 

Cleveland, Ohio 44135 
Introduction

Various new technologies currently under development may enable controlled blade shape 

variability, or so-called blade morphing, to be practically employed in aircraft engine fans and 

compressors in the foreseeable future. Information regarding the potential aerodynamic benefits of 

blade morphing would be useful for ongoing advocacy, prioritizing, and planning of near-term 

research in this area.

The current study is a relatively brief, preliminary computational fluid dynamics (CFD) 

investigation aimed at partially demonstrating and quantifying the aerodynamic potential of fan 

rotor blade morphing. The investigation is intended to provide information useful for near-term 

planning, as well as CFD solution data sets that can be subsequently analyzed using advanced 

acoustic diagnostic tools, if desired, for the purpose of making fan noise comparisons.

Two existing fan system models serve as baselines for the investigation. The first system, 

the Advanced Ducted Propulsor (ADP) fan, version 1, was designed by United Technologies 

Corporation, Pratt & Whitney [1,2]Ü with NASA funding under the Advanced Subsonic 

Technology program. The ADP fan is a 22-inch sub-scale, low-noise research fan/nacelle model 

that has undergone extensive experimental testing in the 9- by 15-foot Low Speed Wind Tunnel 

(LSWT) at the NASA Glenn Research Center [3]. The second system, the Source Diagnostic Test 

(SDT) fan, was designed by the General Electric Corporation with partial funding from NASA 

under the Advanced Subsonic Technology program. The SDT fan is also a 22-inch sub-scale, lownoise research fan/nacelle model that has undergone extensive experimental testing in the LSWT 

[4,5].

Due to limited time and resource constraints, the present study is restricted to fan rotor 

blade morphing only, and a fairly simple blade morphing technique is used. Specifically, spanwiselinear variations in rotor blade-section setting angle are applied to alter the blade shape; that is, the 

blade is linearly retwisted from hub to tip. The amount of morphing is quantified by the ìtip morph 

angleî, which is the change in blade tip-section setting angle relative to the hub-section setting 

angle.

Aerodynamic performance comparisons between morphed-blade and corresponding 

baseline configurations are made on the basis of equal fan system thrust, where rotor rotational 

speed for the morphed-blade fan is varied to change the thrust level for that configuration. Linear 

interpolation or extrapolation of the computed solutions is used to exactly match the baseline 

thrust.

Three simulated flight conditions are included in the investigation. They are takeoff, cruise, 

and approach, where the approach condition is considered here only for the SDT fan system. 

Computationally, flight conditions are set by far-field boundary values, especially Mach number 

and the related static-to-total pressure ratio. Takeoff and approach far-field conditions are the same 

and correspond to standard-day sea-level (SSL) flight at Mach number 0.100. Cruise conditions 

correspond to standard-day flight at 30,000 feet and Mach number 0.800 for the ADP fan system, 

or Mach number 0.780 for the SDT fan system.

Ü Numbers in square brackets indicate references.

NASA/CR—2012-217815 2
Fan System Models

Both fan system models employed for the CFD investigation are fairly complete, including 

all major fan system components such as engine inlet, fan rotor, fan exit guide vane (FEGV) row, 

fan nozzle, and engine nacelle. Of the two systems, the SDT model is simpler because it involves 

only a single flow path along with a fixed-pitch rotor, while the ADP model involves a flow path 

split (into core and bypass streams) between the rotor and FEGV row, and the rotor is variable 

pitch. Both systems are described in more detail below, along with the rotor morphed-blade 

configurations investigated for each.

ADP Fan System

A meridional-plane drawing of the ADP fan system is shown in Figure 1, where all major 

components are depicted and shown to scale. The rotor blade stacking line is the zero-reference 

axial location, and the number of blades for each blade row is indicated in parenthesis. It should be 

noted that downstream of the core inlet guide vane (CIGV) row, the fan system hardware involves 

a long core-duct (not shown) that extends downstream to well beyond the fan nozzle exit. This core 

duct is not included in the CFD model.

Because the ADP rotor is a variable-pitch rotor, there are essentially two reference, or 

baseline, blade shapes for it, one for cruise and one for takeoff. The geometry of the blade itself is 

the same at both conditions, but the variable-pitch blade setting angle is different by 9.0 degrees. 

At cruise, which is the aerodynamic design point, the blades are staggered to be more open; that is, 

they are set to be more aligned with the axial direction. For takeoff the variable-pitch blades are 

restaggered by rotation about the rotor stacking axis to a more closed setting angle. CorrectedÜ

rotational speed also differs for the two operating points: at takeoff the speed is 8750 rpm (100 

percent), and at cruise the speed is 8397 rpm (96.0 percent)á. The corresponding linear tip speeds 

for takeoff and cruise are 840 and 806 ft/sec, respectively. A uniform rotor tip clearance of 0.020 

inches was used for all CFD simulations at both speeds.

Two different blade morphing schedules, or schemes, were applied to the ADP fan rotor, 

with the intent to determine if morphing could be an acceptable substitute for rotor blade reset. It 

was anticipated that this might demonstrate, quantitatively, the potential aerodynamic effectiveness 

of blade morphing, and possibly also show a practical use for it, assuming that the hardware 

mechanisms needed to achieve morphing are substantially lighter and less complex than those 

associated with variable pitch (blade reset).

The first morphing scheme involves an intermediate hub-section setting angle that is 

midway between the baseline takeoff and baseline cruise setting angles, that is, a -4.5 degree hub 

reset relative to the takeoff positionß, with blade tip morph angles of +5.0 and -7.0 degrees at 

Ü The term ìcorrectedî as applied to rotational speed, mass flow rate, or thrust is terse for ìcorrected to standardday sea-level conditionsî.

á For this investigation the corrected rotational speed at takeoff is used as the reference speed, even though the fan 

design point is at cruise.

ß Negative setting-angle changes, including blade tip morph values, correspond to decreased blade section setting 

angles, where the blade chord is oriented more in the axial direction and the blade passages are more open.

NASA/CR—2012-217815 3
takeoff and cruise, respectively. Initially, a tip morph angle of -5.0 degrees was selected for cruise, 

but using an angle of -7.0 degrees provides a significant improvement in aerodynamic 

performance.

The second schedule involves the takeoff hub-section setting angle, corresponding to a 0.0 

degree reset, with blade tip morph angles of 0.0 and -10.0 degrees at takeoff and cruise, 

respectively. Note that since the blade is not morphed at takeoff, the fan operating point there is the 

same as the baseline.

SDT Fan System

A meridional-plane drawing of the SDT fan system is shown in Figure 2. As in the previous 

figure, all major components are depicted and shown to scale, the rotor blade stacking line is the 

zero-reference axial location, and the number of blades for each blade row is indicated in 

parenthesis. Like for the ADP fan, the SDT aerodynamic design point is at the cruise operating 

condition. The SDT rotor blades, however, are fixed at the hub, which is vastly more conventional, 

and the corrected rotational speed of 12,657 rpm (100 percent) for cruise is also used at takeoff. 

The corresponding linear tip speed is 1,215 ft/sec. A uniform rotor tip clearance of 0.034 inches 

was used for all CFD simulations at takeoff and cruise, while a clearance of 0.043 inches was used 

for approach.

Only a single blade morphing schedule was applied to the SDT fan rotor, with the objective 

of improving fan performance. Note that this goal is very different from the one mentioned above 

for the ADP fan. Since the SDT fan aerodynamic design point corresponds to cruise operation, the 

blade geometry there was taken as the reference shape, and morphing was applied only at the 

takeoff and approach operating conditions, where for both a tip morph angle of -5.0 degrees was 

investigated. At approach, with a baseline fan rotational speed of 7,809 rpm (61.7 percent), a 

second tip morph angle of -10.0 degrees was also investigated. Note that with regard to fan 

performance at either of these operating conditions, it should be recognized that there are multiple 

figures of merit; for example, adiabatic efficiency, stall margin, transient response, and noise.

NASA/CR—2012-217815 4
Computational Fluid Dynamics

Flow field simulations were done using two different CFD codes: an axisymmetric viscous 

solver called AVCS, and a three-dimensional viscous turbomachinery solver called TSWIFT. 

Multiple solution domains (grid blocks) were used, with axisymmetric solutions coupled to threedimensional solutions at mixing planes by means of a separate computer program called SMPI, 

developed as a companion program for AVCS and TSWIFT. SMPI was also used to couple rotating 

and stationary three-dimensional solutions together at mixing planes. In general, the threedimensional TSWIFT solver was used for computational domains in and near blade rows, and the 

axisymmetric AVCS solver was used for computational domains sufficiently far away from blade 

rows; that is, far enough away that pitchwise (or circumferential) variations can be ignored. All 

program-to-program data exchange, for mixing planes and direct block-to-block interfaces, was 

performed using a message passing interface called SYNCEXÜ.

The AVCS and TSWIFT codes use similar numerical algorithms; both solve the Reynoldsaveraged Navier-Stokes equations on body-fitted grids using an explicit, finite-difference scheme. 

The codes include viscous terms in all grid directions, except that TSWIFT neglects the mixedderivative viscous terms containing derivatives in the predominantly streamwise grid direction. 

The discretized equations are solved with a multi-stage Runge-Kutta time-marching scheme using 

a spatially varying time step, implicit residual smoothing, and preconditioning [6-9]. All 

simulations discussed herein were run using a 2-stage Runge-Kutta scheme with a CFL number of 

2.5, and using the AUSM+ upwind scheme [10].

The TSWIFT code was derived from, and has the same basic features as, the SWIFT code 

[11] developed by Chima at the NASA Glenn Research Center. TSWIFT also has a fairly general 

multiblock capability (when used with SYNCEX), includes the two-equation SST turbulence 

model developed by Menter [12], and implements Giles' two-dimensional, steady-state, nonreflecting boundary conditions [13-14] at flow inlet, exit, and mixing-plane boundariesá

. Note that 

when a two-equation turbulence model is used, either the Wilcox k- model [15] or the Menter 

SST model, it is necessary to pitchwise average the computed turbulence properties on the 

upstream side of the mixing plane. In that case the turbulence kinetic energy, k, and the ratio of 

turbulence kinetic energy to turbulence dissipation rate, k/, are each mass-averaged, and the 

resulting average values of k and (indirectly)  are used as inflow boundary values for the domain 

on the downstream side of the mixing plane.

Ü SYNCEX (pronounced ìsink-exî) enables two or more executing programs to efficiently exchange data on a 

single computer and/or over a network.

á The SMPI code also implements Giles' two-dimensional, steady-state, non-reflecting boundary conditions.

NASA/CR—2012-217815 5
Computational Grids

Two-dimensional grid blocks for axisymmetric computational domains were generated 

using an algebraic method. In most cases the mesh was restretched using a Poisson partial 

differential equation (PDE) solver, otherwise known as an elliptic grid generator, producing grids 

with better boundary-normal node clustering and spacings, and generally good local orthogonality.

Three-dimensional grid blocks for blade row computational domains, except for rotor 

downstream extension H-grid blocks, were generated using a computer program called TTGRID, 

which is a modified version of TCGRID [16], a three-dimensional grid generator for 

turbomachinery developed by Chima at the NASA Glenn Research Center. TTGRID applies an 

elliptic PDE solver to the streamwise-pitchwise (blade-to-blade) mesh surfaces of blade row Cand H-grids.

The AVCS and TSWIFT codes always directly compute the viscous sublayer solution for 

turbulent boundary layers on solid walls; that is, wall functions are not used. It is important, 

therefore, that node spacings at solid walls be sufficiently small. At all hub and tip endwalls, a 

wall-normal nominal spacing of 0.0001 to 0.0003 inches, depending on location, was specified for 

grid generation, and at all blade/vane surfaces a wall-normal nominal spacing of 0.0002 inches was 

specified. Corresponding inner-variable wall distances, yw

+

, typically range between 1.0 and 3.0 at 

endwall surfaces, and 1.0 and 2.0 at blade surfaces.

More details with figures showing the various grid blocks for the ADP and SDT fan system 

computational domains, respectively, can be found in References 17 and 18. The grid blocks 

reported in these references are not identical to those used in this study, but they are generally very 

similar, with the exception of the ADP rotor blade grid block near the hub endwall. For the current 

investigation the blade is fully attached to the hub, while in the reference the blade-to-hub 

attachment involves a mid-chord trunnion with fore and aft clearance gaps. Grid block sizes for the 

ADP and SDT fan systems are summarized below in Tables 1 and 2, respectively.

NASA/CR—2012-217815 6
Table 1: Computational Grid Blocks for ADP Fan System

Grid Block Size (I × J × K) Nodes

Inlet H-Grid 169 × 1 × 85 14,365

Nozzle H-Grid 169 × 1 × 85 14,365

Nacelle H-Grid 297 × 1 × 65 19,305

Far Field H-Grid 169 × 1 × 45 7,605

Rotor Upstream H-Grid 21 × 26 × 85 46,410

Rotor Blade C-Grid 193 × 49 × 85 803,845

Rotor Tip Clearance O-Grid 161 × 13 × 13 27,209

Rotor Downstream Core H-Grid 49 × 89 × 53 231,133

Rotor Downstream Bypass H-Grid 69 × 89 × 73 448,293

Core Inlet Guide Vane C-Grid 177 × 41 × 53 384,621

Fan Exit Guide Vane C-grid 193 × 45 × 73 634,005

Total All Grid Blocks 2,631,156

Table 2: Computational Grid Blocks for SDT Fan System

Grid Block Size (I × J × K) Nodes

Inlet H-Grid 169 × 1 × 85 14,365

Nozzle H-Grid 169 × 1 × 85 14,365

Nacelle H-Grid 321 × 1 × 65 20,865

Far Field H-Grid 177 × 1 × 45 7,965

Rotor Upstream H-Grid 33 × 42 × 85 117,810

Rotor Blade C-Grid 217 × 49 × 85 903,805

Rotor Tip Clearance O-Grid 179 × 13 × 17 39,559

Rotor Downstream H-Grid 89 × 89 × 81 641,601

Fan Exit Guide Vane C-grid 209 × 45 × 73 686,565

Total All Grid Blocks 2,446,900

NASA/CR—2012-217815 7
Aerodynamic Simulations

All aerodynamic simulations were performed with air modeled as a perfect gas, and with 

the ratio of specific heats, , equal to 1.400. Far-field total/stagnation properties are normalized to 

SSL conditions, so all computed and post-processed solution results directly represent corrected 

values. The effects of flight Mach number and altitude are embedded in the unit Reynolds number 

specified for the computations. At the ADP cruise condition, corresponding to standard day flight 

at 30,000 feet and Mach 0.800, the unit Reynolds numberÜ is 3.047E5 inches-1. Similarly, for 

standard day flight at 30,000 feet and Mach 0.780, which is the SDT cruise condition, the unit 

Reynolds number is 3.008E5 inches-1. Note that at the takeoff and approach conditions, 

corresponding to SSL flight at Mach 0.100, the effect of Mach number was ignored, so the SSL 

unit Reynolds number of 5.915E-05 inches-1 was specifiedá.

The effects of turbulence were modeled using the Menter SST turbulence model [12], with 

freestream turbulence at the upstream far-field computational boundary specified to be 0.5 percent, 

and turbulence (eddy) viscosity set equal to 2.0 times the molecular viscosity. The corresponding 

turbulence kinetic energy depends on the far-field Mach number. For the conditions at 

takeoff/approach, ADP cruise, and SDT cruise, respectively, the calculated values are as follows: 

3.74E-07, 2.13E-05, and 2.03E-05 (dimensionless; multiply by the far-field stagnation speed-ofsound-squared to obtain a dimensional value).

Since the core duct downstream of the CIGV row in the ADP fan system was not modeled 

for the CFD simulations, it was necessary to specify static pressures at the exit boundary of the 

CIGV computational domain. This was done by setting the hub static pressure at a specified value, 

and using simple radial-equilibrium to compute the static pressure distribution from hub to tip. For 

the baseline takeoff and cruise operating points, hub static pressure was adjusted iteratively to 

achieve the desired system bypass ratio (BPR) for each operating point. For the corresponding 

morphed-blade cases however, the hub static pressure from the baseline case was used, so BPR 

varies somewhat between related cases.

Aerodynamic performance comparisons between baseline and corresponding morphedblade configurations are made, as already mentioned, on the basis of equal fan system thrustß

. This 

involves varying fan rotational speed for the morphed blade configuration, thereby changing the 

thrust level for that system. It is possible to iteratively vary the rotational speed so as to ultimately 

match the baseline thrust, but it is normally much faster and more computationally efficient to 

simulate only two cases having thrust levels near the baseline value, and then use linear 

Ü Unit Reynolds number is based on far-field total/stagnation properties and has dimensions of inverse-length. The 

length unit must be the same as that of the computational grid(s).

á At 30,000 feet and flight Mach 0.800, the total-pressure correction factor, , is 0.4529, and the total-temperature 

correction factor, , is 0.8955. At 30,000 feet and flight Mach 0.780, the total-pressure correction factor is 

0.4441, and the total-temperature correction factor is 0.8904. At sea level and flight Mach 0.100, the totalpressure correction factor is 1.0006, and the total-temperature correction factor is 1.0020.

ß Thrust is calculated by integrating flow field properties on the boundaries of a control volume surrounding the 

fan system. Solid surface boundaries and flow boundaries are involved. Control volume boundaries located 

outside the computational domain contribute zero to the net thrust of the system.

NASA/CR—2012-217815 8
interpolation or extrapolation to calculate the thrust-matched quantities. As would be expected, 

most of the resulting fan performance quantities such as mass flow rate, total-pressure ratio, and 

power, differ from corresponding baseline values, but the magnitudes are generally similar.

Quantitative performance comparisons are made using overall stage adiabatic efficiency 

since it reflects most of the significant performance-related differences inherent in the computed 

aerodynamic solutions. Regarding this, it should be noted that for the ADP fan system, stage 

efficiency applies only to the bypass stream. Also note that all adiabatic efficiency values as 

presented herein are relative to conditions upstream of the fan inlet, which means that fan inlet 

total-pressure losses are included in the fan efficiencies.

Two other important performance-related figures of merit for comparing the computed 

aerodynamic solutions are fan stall margin and noise. Although a quantitative assessment and 

comparison of these parameters is well beyond the scope of this investigation, some relevant 

qualitative discussions are provided below where appropriate.

The computational solutions for the ADP and SDT fans systems are presented and 

discussed next. For convenience, Table 3 summarizes the flight conditions considered for the two 

systems. Before proceeding, however, a couple of comments concerning the selected flight Mach 

numbers are appropriate. The first concerns the two cruise Mach numbers, both of which were 

originally specified to be 0.800. While initial CFD solutions showed this value to be good for the 

ADP system, it proved to be excessive for the SDT system, where shock waves generated on the 

outer nacelle surface were fairly strong and noticeably adverse to the boundary layer there. 

Consequently, the flight Mach number for the SDT system was reduce to Mach 0.780. The second 

comment concerns the takeoff and approach Mach number of 0.100. This value is related to 

experimental testing in the 9- by 15-foot Low Speed Wind Tunnel at the NASA Glenn Research 

Center, rather than to aircraft mission profile information.

Table 3: Summary of Flight Conditions for Aerodynamic Simulations

ADP Fan SDT Fan

Cruise

Mach 0.800 @ 30,000 ft

Cruise

Mach 0.780 @ 30,000 ft

Takeoff

Mach 0.100 @ Sea Level

Takeoff

Mach 0.100 @ Sea Level

óóó

Approach

Mach 0.100 @ Sea Level

NASA/CR—2012-217815 9
ADP Fan System at Cruise

Mach number contours for the computed flow field of the ADP baseline fan system at the 

cruise operating condition are shown in Figure 3, where the three-dimensional blade-row solutions 

have been mixed-out averaged in the pitchwise direction. Note that about half of the outer far-field 

domain has been cropped in the upper image, so the outer boundary is much farther away than 

might be inferred from the figure. Only the baseline flow field is shown since it also closely 

represents that of the morphed-blade configurations as wellÜ.

Table 4 below summarizes the different configurations computationally investigated for the 

ADP fan at the cruise condition. Included in the table are corrected fan mass flow rates and system 

bypass ratios. Corresponding fan rotor and fan stage performance quantitiesá are shown in Table 5. 

In each table the red entries indicate values that are linearly extrapolated, based on system thrust, 

from the computed values. Some of the data in Table 5 has been graphed, as presented in Figure 4, 

where the bottom graph shows system thrust, and the top graph shows stage adiabatic efficiency, 

each versus percent rotational speed.

Table 4: ADP Fan Configurations at Cruise (Flight Mach 0.800)

Configurationß

Hub Reset 

Angle 

(degrees)

Tip Morph 

Angle 

(degrees)

Rotational 

Speed 

(percent)

Mass Flow 

Rate 

(lbm/sec)

Bypass

Ratio

Baseline -9.0 0.0 95.96 93.71 13.30

H-4.5 // M-05

-4.5 -5.0 99.00 93.70 13.88

-4.5 -5.0 100.00 93.99 13.85

-4.5 -5.0 98.22 93.47 13.90

H-4.5 // M-07 -4.5 -7.0 96.00 93.45 13.98

H-0.0 // M-10

0.0 -10.0 102.00 93.21 14.56

0.0 -10.0 104.00 93.64 14.57

0.0 -10.0 101.12 93.02 14.56

Ü The qualifications stated here apply to all other results of the same type, presented subsequently.

á Throughout this report, overall aerodynamic quantities are based on spanwise entropy-averaged flow field 

properties (an entropy-conserving method), except for system thrust, which is based on mixed-out-averaged flow 

field properties. Average spanwise quantities are based on pitchwise mixed-out-averaged flow field properties. 

Fan total-pressure ratio and adiabatic efficiency are relative to the inlet freestream condition, and therefore 

include total-pressure losses for the fan/engine inlet.

ß In all configuration labels for the ADP fan, the number following ìHî indicates the hub reset angle, in degrees, 

and the number following ìMî indicates the tip morph angle, in degrees.

NASA/CR—2012-217815 10
Table 5: ADP Fan System Performance at Cruise (Flight Mach 0.800)

Configuration

Rotational 

Speed 

(percent)

Rotor 

Pressure 

Ratio

Rotor 

Adiabatic 

Efficiency

Fan Stage 

Pressure 

Ratio

Fan Stage 

Adiabatic 

Efficiency

Fan System 

Corrected 

Thrust (lbf)

Baseline 95.96 1.2822 0.9218 1.2826 0.8881 615.01

H-4.5 // M-05

99.00 1.2865 0.9084 1.2880 0.8760 624.92

100.00 1.2911 0.9035 1.2924 0.8707 637.57

98.22 1.2829 0.9122 1.2846 0.8801 615.01

H-4.5 // M-07 96.00 1.2836 0.9237 1.2853 0.8909 615.38

H-0.0 // M-10

102.00 1.2861 0.8896 1.2875 0.8578 622.95

104.00 1.2932 0.8782 1.2945 0.8465 640.95

101.12 1.2830 0.8946 1.2844 0.8628 615.01

In Table 4, the BPR of 13.30 for the baseline configuration was taken from the original 

design report [1], and so was predetermined. Information in the design report also gives a designpoint corrected mass flow rate of 91.8 lbm/sec at a stage total-pressure ratio of 1.294. The mass 

flow rate for the CFD solution is 2.1 percent higher, with a lower stage total-pressure ratio of 

1.283. Since the design-point flight Mach number is unknown to the author, it is very possible that 

this performance difference is due largely to a flight Mach number differenceÜ.

The efficiency graph in Figure 4 shows an almost linear relationship between fan stage 

efficiency and rotational speed. This is not a general characteristic, however, but is particular to 

this fan at this operating condition. It appears to be driven primarily by rotor near-tip total-pressure 

losses, where a shock/boundary-layer interaction intensifies as rotational speed increases. 

Corresponding spanwise distributions of rotor adiabatic efficiency are shown in Figure 5, which 

also includes a graph of rotor total-pressure ratio. As can be seen in Figure 5, there is a clear trend 

of decreasing efficiency near the tip as rotational speed increases. The shock/boundary-layer 

interaction producing this trend is shown in Figure 6, which shows Mach number contours for the 

rotor flow field at 85 percent of span, for two of the configurations simulated; namely, the baseline 

configuration and the -10 degree morphed-blade configuration. Examination of the two flow fields 

reveals that the suction surface boundary layer remains attached downstream of the shock wave for 

the baseline case, while for the morphed-blade case a substantial separation exists.

Referring again to Figure 4, as well as Table 5, it can be seen that the morphed-blade 

configuration involving a -4.5 degree hub reset and a -7.0 degree tip morph is best. This 

configuration has the same rotational speed as the baseline, with a slightly higher fan stage 

adiabatic efficiency (+0.003 for the bypass stream). Only near the hub, locally, is the efficiency 

substantially below that of the baseline configuration (see Figure 5). The worst configuration is the 

Ü Table XII in Reference 1 indicates an inlet corrected flow rate of 97.2 lbm/sec at cruise Mach number 0.82.

NASA/CR—2012-217815 11
one with a 0.0 degree hub reset and a -10.0 degree tip morph. From Figure 4 it seems reasonable to 

infer that the efficiency could be improved, that is for a configuration with a 0.0 degree hub reset, 

by increasing the magnitude of the tip morph angle, thereby lowering rotational speed. However, in 

order to match the baseline speed and efficiency, a tip morph angle of somewhere around -16.0 

degrees would be needed, which is probably an impractical amount of blade morphing.

ADP Fan System at Takeoff

Mach number contours for the computed flow field of the ADP baseline fan system at the 

takeoff operating condition are shown in Figure 7. The fan mass flow rate for the computed 

solution is 78.50 lbm/sec at a BPR of 13.50, which agrees reasonably well with the measured flow 

rate of 79.20 lbm/sec at a BPR of 11.50. Adjusting the computed flow rate to roughly account for 

the difference in BPR, by assuming a fixed bypass flow, gives a fan mass flow rate of 79.44 

lbm/sec, which is in closer agreement. Note that the higher BPR of 13.50 was selected for the 

present investigation because it is more consistent with the cruise/design value, and also because 

higher BPR conditions tend to make the fan system performance less sensitive to small BPR 

changes, such as occurs with morphed-blade configurations.

For the takeoff condition only a single ADP morphed-blade configuration was investigated. 

The fan configuration and performance data are summarized below in Tables 6 and 7, respectively. 

As earlier, red entries indicate values that have been linearly extrapolated, based on system thrust, 

from computed values. Notice that the hub reset angle of -4.5 degrees is the same as that 

investigated for cruise, the intent being that tip morph angle varies from -7.0 degrees at takeoff to 

+5.0 degrees at cruise. Graphs of key performance data from Table 7 are presented in Figure 8, the 

bottom graph showing system thrust, and the top showing stage adiabatic efficiency, each versus 

percent rotational speed. Related spanwise profiles of rotor total-pressure ratio and adiabatic 

efficiency are provided in Figure 9.

All the results for takeoff show the morphed-blade aerodynamic performance to be nearly 

the same as the baseline performance. The flat trend in adiabatic efficiency (see Figures 8 and 9) as 

rotational speed varies is quite different from the trend at cruise, and is due to the lack of any 

sensitive fluid dynamic mechanisms that would cause substantial changes in total-pressure loss. 

There is, however, a possible negative performance impact for the morphed-blade configuration; 

namely, a potential reduction in stall margin, depending on where rotor stall initiates.

Table 6: ADP Fan Configurations at Takeoff (Flight Mach 0.100)

Configuration

Hub Reset 

Angle 

(degrees)

Tip Morph 

Angle 

(degrees)

Rotational 

Speed 

(percent)

Mass Flow 

Rate 

(lbm/sec)

Bypass

Ratio

Baseline 0.0 0.0 100.00 78.50 13.50

H-4.5 // M+05

-4.5 5.0 96.00 76.75 12.89

-4.5 5.0 98.00 78.32 12.96

-4.5 5.0 98.51 78.72 12.98

NASA/CR—2012-217815 12
Table 7: ADP Fan System Performance at Takeoff (Flight Mach 0.100)

Configuration

Rotational 

Speed 

(percent)

Rotor 

Pressure 

Ratio

Rotor 

Adiabatic 

Efficiency

Fan Stage 

Pressure 

Ratio

Fan Stage 

Adiabatic 

Efficiency

Fan System 

Corrected 

Thrust (lbf)

Baseline 100.0 1.2808 0.9450 1.2847 0.9194 1336.8

H-4.5 // M+05

96.0 1.2631 0.9464 1.2660 0.9203 1263.7

98.0 1.2755 0.9460 1.2785 0.9200 1321.9

98.5 1.2787 0.9459 1.2817 0.9199 1336.8

Although not shown, for the takeoff condition the morphed blade has higher leading-edge 

flow incidence angles over much of the span, especially near the hub, and to a lesser extent at 

midspan. In Figure 10, blade surface pressure coefficient distributions for the morphed and 

baseline rotor blades at 32 percent of span are compared. The morphed-blade distribution shows 

greater forward aerodynamic blade loading (vertical distance between suction- and pressuresurface static pressures), which is indicative of the higher incidence angles and would be expected 

to correspond to less stall tolerance for that particular region of the blade span. It must be 

emphasized, however, that this may or may not be significant, depending on where along the span 

the rotor tends to stall initially. Furthermore, whether or not a decrease in stall margin is important 

or critical depends on available and required stall margin levels, both of which are beyond the 

scope of this study.

SDT Baseline Fan System at Cruise and Takeoff

Mach number contours for the computed flow field of the SDT baseline fan system at the 

cruise operating condition are shown in Figure 11, and similarly, contours for the takeoff operating 

condition are shown in Figure 12. No morphed-blade configurations were investigated for the 

cruise condition, which corresponds to the fan aerodynamic design point, but the computed 

solution provides baseline performance results that are useful as a reference. Also, comparisons 

between the cruise and takeoff results show how the baseline fan performance varies between the 

two operating points.

Computed results for the cruise condition can only be compared to design values because 

experimental testing has not been performed at that condition. From the CFD simulation for cruise, 

the corrected fan mass flow rate is 100.9 lbm/sec, with a corresponding stage total-pressure ratio of 

1.451 and a stage adiabatic efficiency of 0.896. These results are in good agreement with reported 

design values [4,5]; specifically, a mass flow rate of 100.5 lbm/sec at a stage pressure ratio of 

1.470. It might be noted that the flight Mach number for the design-point cruise condition is not 

exactly known, so it may differ somewhat from that used for the simulation. In view of the 

computed transonic flow over the outer nacelle surface (see Figure 11), it seems unlikely that the 

flight Mach number would be larger than the value of 0.780 selected for this study. A slightly lower 

Mach number would only bring the computed and design values into closer agreement.

NASA/CR—2012-217815 13
For the takeoff condition, the computational results can be directly compared to 

experimental measurement values [5]. The agreement is generally good, although there are some 

notable inconsistencies. In particular, the computed and measured fan mass flow rates are in very 

good agreement, both being the same at 97.2 lbm/sec. Fan stage total-pressure ratio, however, 

differs significantly, at 1.483 for the computed solution, and 1.509 from the measurements. 

Corresponding stage adiabatic efficiencies are 0.894 and 0.914, respectively. This apparent 

difference in stage pressure ratio is inconsistent with the agreement in mass flow rate, unless there 

is an error in the fan nozzle geometry model, which seems very unlikely considering the good 

agreement between CFD and design values at cruise. In any case, efforts to resolve and explain 

these performance discrepancies are beyond the scope of this investigation and would require 

detailed comparisons of the computational and available measurement data.

Spanwise distributions of SDT baseline rotor total-pressure ratio and adiabatic efficiency at 

the cruise and takeoff operating conditions are compared in Figure 13. From the total-pressure 

distributions it can be seen that rotor aerodynamic loading increases over only the outer threefourths of the span, but especially at the tip, as flow rate decreases; that is, as fan pressure ratio 

increases from cruise to takeoff. Accompanying the increase in aerodynamic loading is a 

corresponding enlargement of the tip clearance secondary flow region where, as indicated by the 

adiabatic efficiency distributions, there is an increase in local total-pressure losses. From this 

performance trend the rotor can be expected to stall at its tip, which is not unusual and is 

significant to the current investigation.

SDT Fan System at Takeoff

Summarized below in Table 8 are the configurations investigated for the SDT fan at the 

takeoff condition, and in Table 9 the corresponding fan rotor and fan stage performance quantities 

are provided. As in previous tables, red entries indicate values that are linearly interpolated, based 

on system thrust, from the computed values. In Figure 14, graphs of the fan system thrust and the 

fan stage adiabatic efficiency, each versus percent rotational speed, are shown.

Table 8: SDT Fan Configurations at Takeoff (Flight Mach 0.100)

ConfigurationÜ

Tip Morph 

Angle 

(degrees)

Rotational 

Speed 

(percent)

Mass Flow 

Rate 

(lbm/sec)

Baseline 0.0 100.00 97.22

M+05

5.0 100.00 93.21

5.0 105.00 97.52

5.0 104.86 97.40

Ü In all configuration labels for the SDT fan, the number following ìMî indicates the tip morph angle, in degrees.

NASA/CR—2012-217815 14
Table 9: SDT Fan System Performance at Takeoff (Flight Mach 0.100)

Configuration

Rotational 

Speed 

(percent)

Rotor 

Pressure 

Ratio

Rotor 

Adiabatic 

Efficiency

Fan Stage 

Pressure 

Ratio

Fan Stage 

Adiabatic 

Efficiency

Fan System 

Corrected 

Thrust (lbf)

Baseline 100.00 1.4956 0.9134 1.4832 0.8935 2258.8

M+05

100.00 1.4443 0.9276 1.4337 0.9081 2051.0

105.00 1.4942 0.9169 1.4822 0.8976 2264.5

104.86 1.4928 0.9172 1.4809 0.8979 2258.9

In terms of overall adiabatic efficiency, the morphed-blade configuration represents a minor 

but significant improvement over the baseline, the difference being about 0.004. When the 

corresponding spanwise variations in rotor adiabatic efficiency and total-pressure ratio, shown in 

Figure 15, are considered, however, the performance differences are much more prominent. Notice 

in particular the change in spanwise aerodynamic loading, as represented by rotor total-pressure 

ratio, where for the morphed-blade configuration the loading at the outer one-third of span has 

been substantially reduced. Not only is the total-pressure ratio lower there, but the rotational speed 

is higher by 5 percent.

From an operational standpoint, this performance difference is potentially a major 

improvement, as it is almost certain to mean a much larger takeoff stall margin for the fan. Recall 

the earlier comment, made when comparing the baseline cruise and takeoff spanwise loading 

distributions, about SDT rotor stall occurring at the tip. Blade-to-blade flow field comparisons 

reinforce this conclusion. In Figure 16, for example, relative Mach number contours for the 

baseline and morphed-blade CFD solutions at 78 percent of span are compared, revealing 

substantially different bow and passage shock wave structures. For the baseline rotor the bow 

shock wave is fully detached and part of a strong ìnormalî shock that is located at the coveredpassage entrance. In contrast, the bow shock wave for the morphed-blade rotor is only locally 

detached, with a weak oblique shock running into the covered passage, and the supersonic flow 

finally terminated by a ìnormalî passage shock located well within the covered passage. 

Consequently, the morphed-blade rotor can be expected to tolerate a much larger increase in 

downstream static pressure, and therefore also in rotor total-pressure ratio, before stall initiates in 

the tip region. Stall at the hub is not expected, because as already shown for this rotor at this 

operating condition, hub loading tends not to increase as overall aerodynamic loading increases 

(see Figure 13).

Another possible advantage of the morphed-blade rotor is related to the upstream-running 

weak oblique shock waves which, as can be seen from Figure 16, are much weaker compared to 

those of the baseline rotor. This might have a significant impact on upstream fan noise at takeoff. 

In the absence of an acoustic analysis, the potential is at least suggested by comparing the static 

pressure fields upstream of the rotor, near the grid boundary, as is done in Figure 17. In the figure, 

the pitchwise variations in static pressure near the rotor tip are easily seen to be greater and more 

extensive for the baseline flow field than for the morphed-blade flow field.

NASA/CR—2012-217815 15
SDT Fan System at Approach

Mach number contours for the computed flow field of the SDT baseline fan system at the 

approach operating condition are shown in Figure 18. The different configurations investigated for 

the SDT fan at approach are summarized below in Table 10, and corresponding fan rotor and fan 

stage performance quantities are presented in Table 11. Again, as in previous tables, red entries 

indicate values that are linearly interpolated, based on system thrust, from computed values. 

Graphs of the fan system thrust and fan stage adiabatic efficiency, each versus percent rotational 

speed, are shown in Figure 19, and in Figure 20, graphs of corresponding spanwise distributions of 

rotor total-pressure ratio and adiabatic efficiency are shown.

Table 10: SDT Fan Configurations at Approach (Flight Mach 0.100)

Configuration

Tip Morph 

Angle 

(degrees)

Rotational 

Speed 

(percent)

Mass Flow 

Rate 

(lbm/sec)

Baseline 0.0 61.70 57.96

M+05

5.0 64.17 57.91

5.0 64.79 58.47

5.0 64.30 58.02

M+10

10.0 67.87 58.43

10.0 67.44 58.04

Table 11: SDT Fan System Performance at Approach (Flight Mach 0.100)

Configuration

Rotational 

Speed 

(percent)

Rotor 

Pressure 

Ratio

Rotor 

Adiabatic 

Efficiency

Fan Stage 

Pressure 

Ratio

Fan Stage 

Adiabatic 

Efficiency

Fan System 

Corrected 

Thrust (lbf)

Baseline 61.70 1.1545 0.8933 1.1502 0.8699 715.20

M+05

64.17 1.1534 0.9138 1.1494 0.8908 711.99

64.79 1.1566 0.9132 1.1526 0.8902 727.82

64.30 1.1540 0.9137 1.1500 0.8907 715.20

M+10

67.87 1.1566 0.9223 1.1525 0.8992 726.20

67.44 1.1544 0.9227 1.1503 0.8996 715.20

NASA/CR—2012-217815 16
As can be seen both in Table 11 and in Figure 19, the blade morphing gives a relatively 

large increase in adiabatic efficiency; specifically, fan efficiency increases by about 0.02 and 0.03, 

respectively, for 5.0 and 10.0 degrees of tip morphing. Comparing in Figure 20 the different 

spanwise distributions of rotor adiabatic efficiency reveals that virtually all of the increase is 

achieved at the outer half of the span. These results are not surprising, and in fact were anticipated, 

because the outer-span blade aerodynamics for the baseline SDT rotor at the approach operating 

point involve high incidence angles, with a leading-edge boundary layer separation on the suction 

side of the blade. Although the boundary layer reattaches, it remains relatively thick. Tip morphing 

reduces the local incidence and aerodynamic loading, and therefore the size and severity of the 

separation, resulting in thinner suction surface boundary layers. To show this, blade-to-blade 

contour plots of relative Mach number for the baseline and 10.0 degree tip-morph rotor solutions at 

75 percent of span are provided in Figure 21. In the figure, the differences in leading-edge flow 

incidence and suction surface boundary thickness are apparent. Note that the flow field solution for 

5.0 degrees of tip morphing is, as might be expected, intermediate between the two solutions 

represented in Figure 21.

Finally, referring again to Figure 20, the spanwise distributions in rotor total-pressure ratio, 

when considered along with the differences in rotational speed, show a dramatic decrease in rotor 

tip aerodynamic loading as the amount of tip morphing increases. In general, larger tip morph 

angles (more blade twist) with higher rotational speeds increase the total-pressure ratio at the hub 

and lower it at the tip, while the overall total-pressure ratio remains about the same. The effect of 

this loading change should be a substantial increase in stall margin at approach, especially in view 

of the decreased tip incidence angles and improved suction surface aerodynamics that accompany 

the reduction in loading.

Concluding Remarks

The results of this preliminary investigation confirm that rotor blade morphing could be a 

useful technology, with the potential to enable significant improvements in fan aerodynamic 

performance. Even though the study is very limited in scope and confined to simple geometric 

perturbations of two existing fan systems, the aerodynamic effectiveness of blade morphing is 

demonstrated by the configurations analyzed.

For the ADP fan system the basic objective was to determine if rotor blade morphing could 

serve as an acceptable aerodynamic replacement for variable-pitch blade restaggering, a feature of 

the baseline design. Two operating conditions were considered, cruise and takeoff. Baseline levels 

of adiabatic efficiency were achieved without difficulty at both conditions, and resulting 

aerodynamic qualities appear to be good. On the negative side, there is a possible reduction in stall 

margin at the takeoff condition, but time and resource constraints precluded further investigation. 

For perspective, it might be added that a variable-pitch rotor, especially one with very good 

aerodynamic performance like the ADP rotor, already has the benefit of ìvariable geometryî, so 

performance improvement through blade morphing was not an objective.

The SDT fan system with its fixed-geometry rotor provided a realistic opportunity for 

aerodynamic improvement. Since the fan aerodynamic design point is at the cruise condition, 

improvements at the takeoff and approach operating points were sought, and to a large extent 

NASA/CR—2012-217815 17
achieved. At takeoff, a relatively small improvement in adiabatic efficiency was realized with 

blade morphing, but a very substantial increase in stall margin could be inferred from the 

computational results, as well as a potential reduction in upstream fan noise. At approach, 

relatively large increases in both adiabatic efficiency and expected stall margin were attained, with 

a substantially thinner boundary layer on the blade suction surface, and a much smaller separation 

bubble at the blade leading edge. The observed improvement in rotor flow quality, especially the 

thinner boundary layers and resulting wakes, might be expected to also possibly reduce broadband 

noise.

Despite the positive results of this investigation, the limitations involved should not be 

underestimated. This work provides only a glimpse of what might be gained through blade 

morphing when incorporated as part of an original fan design, where the fan must be integrated and 

optimized with the engine and aircraft. Even the aircraft mission profile could be affected since it 

is closely coupled to propulsion system performance over a wide range of operating conditions. 

From this perspective it seems possible, at least conceptually, that blade morphing might yield 

unanticipated benefits and/or lead to unexpected performance capabilities. If consideration is 

limited only to the fan itself, even then it seems likely that an original design incorporating blade 

morphing would better capitalize on the variability that morphing offers; that is, compared to the 

results from this limited investigation.

Finally, it is important to note that the selection of particular configurations for the present 

investigation implies nothing about the relative merits of the selected configurations versus other 

possible morphing schemes and/or applications. For example, the benefits of stator blade (or guide 

vane) morphing, compressor blade morphing, spanwise non-linear morphing, and blade camber 

morphing, to name a few, were not considered or studied in this work. Also, a less conventional, 

but potentially suitable application for blade morphing might be supersonic aircraft propulsion, 

where engine cycle variability is normally much greater than that needed for subsonic propulsion.

References

1. Hobbs, David E., Neubert, Robert J., Malmborg, Eric W., Philbrick, Daniel H., Spear, 

David A., ìLow Noise Research Fan Stage Design,î NASA Contractor Report (NASA 

internal use only; contains Pratt & Whitney proprietary data), March 1995.

2. Hobbs, David E., Neubert, Robert J., Malmborg, Eric W., Philbrick, Daniel H., Spear, 

David A., ìLow Noise Research Fan Stage Design,î NASA CR-195382, March 1995.

3. Jeracki, Robert J., ìComprehensive Report of Fan Performance From Duct Rake 

Instrumentation on 1.294 Pressure Ratio, 806 ft/sec Tip Speed Turbofan Simulator 

Models,î NASA TM-2006-213863, February 2006.

4. Podboy, Gary G., Krupar, Martin J., Hughes, Christopher E., Woodward, Richard P., ìFan 

Noise Source Diagnostic Test  LDV Measured Flow Field Results,î NASA TM-2003-

212330 (also AIAA-2002-2431), July 2003.

5. Hughes, Christopher E., Jeracki, Robert J., Woodward, Richard P., ìFan Noise Source 

Diagnostic Test  Rotor Alone Aerodynamic Performance Results,î NASA TM-2005-

211681 (also AIAA-2002-2426), April 2005.

NASA/CR—2012-217815 18
6. Chima, R. V., ìViscous Three-Dimensional Calculations of Transonic Fan Performance,î in 

CFD Techniques for Propulsion Applications, AGARD Conference Proceedings No. CP510, AGARD, Neuilly-Sur-Seine, France, February 1992, pp. 21-1 to 21-19 (also NASA 

TM-103800).

7. Chima, R. V., and Yokota, J. W., ìNumerical Analysis of Three-Dimensional Viscous Flows 

in Turbomachinery,î AIAA Journal, Vol. 28, No. 5, May 1990, pp. 798-806.

8. Tweedt, Daniel L., and Chima, Rodrick V., ìRapid Numerical Simulation of Viscous 

Axisymmetric Flow Fields,î NASA TM-107103 (also AIAA-96-0449), November 1995.

9. Tweedt, Daniel L., Chima, Rodrick V., and Turkel, Eli, ìPreconditioning for Numerical 

Simulation of Low Mach Number Three-Dimensional Viscous Turbomachinery Flows,î 

NASA TM-113120 (also ICOMP-97-11 and AIAA-97-1828), October 1997.

10. Chima, R. V. and Liou, M.-S., ìComparison of the AUSM+ and H-CUSP Schemes for 

Turbomachinery Applications,î AIAA Paper AIAA-2003-4120 (also NASA TM-2003-

212457), June 2003.

11. Chima, R. V., ìSwift ñ Multiblock Analysis Code for Turbomachinery, User's Manual and 

Documentation,î Version 300, August 2003.

12. Menter, R. F., ìTwo-Equation Eddy-Viscosity Turbulence Models for Engineering 

Applications,î AIAA Journal, Vol. 32, No. 8, August 1994, pp. 1598-1605 (also NASA 

TM-111958).

13. Giles, Michael B., ìNonreflecting Boundary Conditions for Euler Equation Calculations,î 

AIAA Journal, Vol. 28, No. 12, December 1990, pp. 2050-2058.

14. Giles, Michael, ìUNSFLO: A Numerical Method For The Calculation Of Unsteady Flow In 

Turbomachinery,î GTL Report No. 205, May 1991, pp. 45-56.

15. Wilcox, David C., Turbulence Modeling for CFD, DCW Industries, Inc., La CaÒada, CA, 

Third Edition, 2006.

16. Chima, R. V., ìTCGRID 3-D Grid Generator for Turbomachinery, User's Manual and 

Documentation,î Version 300, July 2003.

17. Tweedt, Daniel L., ìComputational Aerodynamic Simulations of an 840 ft/sec Tip Speed, 

Advanced Ducted Propulsor Fan System Model for Acoustic Methods Assessment and 

Development,î NASA CR to be published, March 2012.

18. Tweedt, Daniel L., ìComputational Aerodynamic Simulations of a 1215 ft/sec Tip Speed 

Transonic Fan System Model for Acoustic Methods Assessment and Development,î NASA 

CR to be published, March 2012.

NASA/CR—2012-217815 19
Figure 1: Meridional-Plane Drawing of ADP Fan System

Figure 2: Meridional-Plane Drawing of SDT Fan System

-10 -5 0 5 10 15 20

Axial Coordinate, x (inches)

0

5

10

15

Radial Coordinate, 

r (inches)

FEGV

(45)

Bypass

Nozzle

Inlet

Nacelle

Spinner

Fan

Rotor

(18)

 CIGV

 (63)

Numbers in Parentheses

are Blade Counts

-10 -5 0 5 10 15 20

Axial Coordinate, x (inches)

0

5

10

15

Radial Coordinate, 

r (inches)

FEGV

(54)

Nozzle

Inlet

Nacelle

Spinner

Fan

Rotor

(22)

Numbers in Parentheses

are Blade Counts

NASA/CR—2012-217815 20
Figure 3: Computed Flow Field for ADP Baseline Fan System at Cruise;

Pitchwise-Averaged Flow; Mach Number Contours

Pitchwise-Averaged Flow Field

ADP Baseline Fan System at Cruise Condition

0.00 0.50 1.00

Mach Number

Flow

NASA/CR—2012-217815 21
Figure 4: Overall Performance of ADP Fan System at Cruise

0.84

0.86

0.88

0.90

Stage Adiabatic Efficiency, 

η

ADP Fan at Cruise Condition

Flight Mach Number 0.800

94 96 98 100 102 104 106

Corrected Rotational Speed, N

c

 (percent)

610

620

630

640

Corrected Thrust, 

Τ (lbf)

CFD Baseline

CFD Morphed

Extrapolated

0.0° hub reset

−10.0° tip morph

−4.5° hub reset

−5.0° tip morph

−4.5° hub reset

−7.0° tip morph

NASA/CR—2012-217815 22
Figure 5: Spanwise Distributions of ADP Rotor Total-Pressure Ratio

and Adiabatic Efficiency at Cruise

1.0 1.1 1.2 1.3 1.4

Total-Pressure Ratio, p t / p tð∞

0.0

0.2

0.4

0.6

0.8

1.0

Fraction of Span from Hub

0.6 0.7 0.8 0.9 1.0

Adiabatic Efficiency, η R

0.0

0.2

0.4

0.6

0.8

1.0

Fraction of Span from Hub

Baseline // Nc = 96%

H-4.5 // M-07 // Nc = 96%

H-4.5 // M-05 // Nc = 99%

H-0.0 // M-10 // Nc = 102%

ADP Rotor at Cruise Condition

NASA/CR—2012-217815 23
Figure 6: Computed Blade-to-Blade Flow Fields for ADP Rotor at Cruise;

85 Percent of Span; Relative Mach Number Contours

85 Percent of Span from Hub

ADP Rotor at Cruise Condition

Baseline

Nc = 96%

Flow

Hub Reset 0.0°

Tip Morph -10.0°

Nc = 102%

0.0

1.0

0.5

Relative

Mach

Number

1.5

Rotation

NASA/CR—2012-217815 24
Figure 7: Computed Flow Field for ADP Baseline Fan System at Takeoff;

Pitchwise-Averaged Flow; Mach Number Contours

0.00 0.50 0.75

Mach Number

0.25

Pitchwise-Averaged Flow Field

ADP Baseline Fan System at Takeoff Condition

Flow

NASA/CR—2012-217815 25
Figure 8: Overall Performance of ADP Fan System at Takeoff

0.90

0.91

0.92

0.93

Stage Adiabatic Efficiency, 

η

ADP Fan at Takeoff Condition

Flight Mach Number 0.100

95 96 97 98 99 100 101

Corrected Rotational Speed, N

c

 (percent)

1260

1280

1300

1320

1340

Corrected Thrust, 

Τ (lbf)

CFD Baseline

CFD Morphed

Extrapolated

−4.5° hub reset

+5.0° tip morph

NASA/CR—2012-217815 26
Figure 9: Spanwise Distributions of ADP Rotor Total-Pressure Ratio

and Adiabatic Efficiency at Takeoff

1.0 1.1 1.2 1.3 1.4

Total-Pressure Ratio, p t / p tð∞

0.0

0.2

0.4

0.6

0.8

1.0

Fraction of Span from Hub

0.80 0.85 0.90 0.95 1.00

Adiabatic Efficiency, η R

0.0

0.2

0.4

0.6

0.8

1.0

Fraction of Span from Hub

Baseline // Nc = 100%

H-4.5 // M+05 // Nc = 98%

ADP Rotor at Takeoff Condition

NASA/CR—2012-217815 27
Figure 10: Blade Surface Pressure Coefficient Distributions for ADP Rotor

at Takeoff; 32 Percent Span

0.0 0.2 0.4 0.6 0.8 1.0

Fraction of Meridional Chord

-1.0

-0.5

0.0

0.5

1.0

Rotor Blade Surface Pressure Coefficient, 

Cp

Baseline // Nc = 100%

H-4.5 // M+05 // Nc = 98%

ADP Rotor at Takeoff Condition

32 Percent of Span from Hub

NASA/CR—2012-217815 28
Figure 11: Computed Flow Field for SDT Baseline Fan System at Cruise;

Pitchwise-Averaged Flow; Mach Number Contours

Pitchwise-Averaged Flow Field

SDT Baseline Fan System at Cruise Condition

0.00 0.50 1.00

Mach Number

Flow

NASA/CR—2012-217815 29
Figure 12: Computed Flow Field for SDT Baseline Fan System at Takeoff;

Pitchwise-Averaged Flow; Mach Number Contours

Pitchwise-Averaged Flow Field

SDT Baseline Fan System at Takeoff Condition

0.00 0.50 0.75

Mach Number

0.25 1.00

Flow

NASA/CR—2012-217815 30
Figure 13: Spanwise Distributions of SDT Baseline Rotor Total-Pressure Ratio

and Adiabatic Efficiency at Cruise and Takeoff

1.3 1.4 1.5 1.6

Total-Pressure Ratio, p t / p tð∞

0.0

0.2

0.4

0.6

0.8

1.0

Fraction of Span from Hub

0.80 0.85 0.90 0.95 1.00

Adiabatic Efficiency, η R

0.0

0.2

0.4

0.6

0.8

1.0

Fraction of Span from Hub

Takeoff Condition

Cruise Condition

SDT Baseline Rotor at Design Rotational Speed

NASA/CR—2012-217815 31
Figure 14: Overall Performance of SDT Fan System at Takeoff

0.88

0.89

0.90

0.91

Stage Adiabatic Efficiency, 

η

SDT Fan at Takeoff Condition

Flight Mach Number 0.100

99 100 101 102 103 104 105 106

Corrected Rotational Speed, N

c

 (percent)

2000

2100

2200

2300

Corrected Thrust, 

Τ (lbf)

CFD Baseline

CFD Morphed

Interpolated

+5.0° tip morph

NASA/CR—2012-217815 32
Figure 15: Spanwise Distributions of SDT Rotor Total-Pressure Ratio

and Adiabatic Efficiency at Takeoff

1.3 1.4 1.5 1.6

Total-Pressure Ratio, p t / p tð∞

0.0

0.2

0.4

0.6

0.8

1.0

Fraction of Span from Hub

0.80 0.85 0.90 0.95 1.00

Adiabatic Efficiency, η R

0.0

0.2

0.4

0.6

0.8

1.0

Fraction of Span from Hub

Baseline // Nc = 100%

M+05 // Nc = 105%

SDT Rotor at Takeoff Condition

NASA/CR—2012-217815 33
Figure 16: Computed Blade-to-Blade Flow Fields for SDT Rotor at Takeoff;

78 Percent of Span; Relative Mach Number Contours

78 Percent of Span from Hub

SDT Rotor at Takeoff Condition

Baseline

Nc = 100%

Tip Morph 5.0°

Nc = 105%

Flow

0.0 0.5 1.0

Relative Mach Number

1.5

Rotation

NASA/CR—2012-217815 34
Figure 17: Computed Static-Pressure Fields Upstream of SDT Rotor at Takeoff

Rotor Upstream Static Pressure Field

SDT Rotor at Takeoff Condition

0.78 0.80 0.82

Dimensionless Static Pressure

0.84

Baseline

Nc = 100%

Tip Morph 5.0°

Nc = 105%

0.86

Rotation

NASA/CR—2012-217815 35
Figure 18: Computed Flow Field for SDT Baseline Fan System at Approach;

Pitchwise-Averaged Flow; Mach Number Contours

Pitchwise-Averaged Flow Field

SDT Baseline Fan System at Approach Condition

0.00 0.25 0.50

Mach Number

Flow

NASA/CR—2012-217815 36
Figure 19: Overall Performance of SDT Fan System at Approach

0.87

0.88

0.89

0.90

Stage Adiabatic Efficiency, 

η

SDT Fan at Approach Condition

Flight Mach Number 0.100

60 62 64 66 68 70

Corrected Rotational Speed, N

c

 (percent)

710

715

720

725

730

Corrected Thrust, 

Τ (lbf)

CFD Baseline

CFD Morphed

Interpolated

+10.0° tip morph

+5.0° tip morph

NASA/CR—2012-217815 37
Figure 20: Spanwise Distributions of SDT Rotor Total-Pressure Ratio

and Adiabatic Efficiency at Approach

1.05 1.10 1.15 1.20

Total-Pressure Ratio, p t / p tð∞

0.0

0.2

0.4

0.6

0.8

1.0

Fraction of Span from Hub

0.7 0.8 0.9 1.0

Adiabatic Efficiency, η R

0.0

0.2

0.4

0.6

0.8

1.0

Fraction of Span from Hub

Baseline // Nc = 61.7%

M+05 // Nc = 64.2%

M+10 // Nc = 67.9%

SDT Rotor at Approach Condition

NASA/CR—2012-217815 38
Figure 21: Computed Blade-to-Blade Flow Fields for SDT Rotor at Approach;

75 Percent of Span; Relative Mach Number Contours

75 Percent of Span from Hub

SDT Rotor at Approach Condition

Baseline

Nc = 61.7%

Tip Morph 10.0°

Nc = 67.9%

0.0 0.5 1.0

Relative Mach Number

Rotation

Flow

NASA/CR—2012-217815 39
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4. TITLE AND SUBTITLE

Preliminary Aerodynamic Investigation of Fan Rotor Blade Morphing 

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NNC06BA07B 

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Tweedt, Daniel, L. 

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NNC10E541 

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WBS 473452.02.03.07.06.01.01 

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13. SUPPLEMENTARY NOTES

14. ABSTRACT

Various new technologies currently under development may enable controlled blade shape variability, or so-called blade morphing, to be practically employed in aircraft 

engine fans and compressors in the foreseeable future. The current study is a relatively brief, preliminary computational fluid dynamics investigation aimed at partially 

demonstrating and quantifying the aerodynamic potential of fan rotor blade morphing. The investigation is intended to provide information useful for near-term planning, as 

well as aerodynamic solution data sets that can be subsequently analyzed using advanced acoustic diagnostic tools, for the purpose of making fan noise comparisons. Two 

existing fan system models serve as baselines for the investigation: the Advanced Ducted Propulsor fan with a design tip speed of 806 ft/sec and a pressure ratio of 1.294, and 

the Source Diagnostic Test fan with a design tip speed of 1215 ft/sec and a pressure ratio of 1.470. Both are 22-in. sub-scale, low-noise research fan/nacelle models that have 

undergone extensive experimental testing in the 9- by 15-foot Low Speed Wind Tunnel at the NASA Glenn Research Center. The study, restricted to fan rotor blade morphing 

only, involves a fairly simple blade morphing technique. Specifically, spanwise-linear variations in rotor blade-section setting angle are applied to alter the blade shape; that is, 

the blade is linearly retwisted from hub to tip. Aerodynamic performance comparisons are made between morphed-blade and corresponding baseline configurations on the 

basis of equal fan system thrust, where rotor rotational speed for the morphed-blade fan is varied to change the thrust level for that configuration. The results of the 

investigation confirm that rotor blade morphing could be a useful technology, with the potential to enable significant improvements in fan aerodynamic performance. Even 

though the study is very limited in scope and confined to simple geometric perturbations of two existing fan systems, the aerodynamic effectiveness of blade morphing is 

demonstrated by the configurations analyzed. In particular, for the Advanced Ducted Propulsor fan it is demonstrated that the performance levels of the original variable-pitch 

baseline design can be achieved using blade morphing instead of variable pitch, and for the Source Diagnostic Test fan the performance at important off-design operating 

points is substantially increased with blade morphing.

15. SUBJECT TERMS

Fan blades; Rotor blades; Torsion; Twisting; Angle of attach 

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