Daniel L. Tweedt

AP Solutions, Inc., Cleveland, Ohio

Computational Aerodynamic Simulations of an

840 ft/sec Tip Speed Advanced Ducted Propulsor 

Fan System Model for Acoustic Methods

Assessment and Development

NASA/CR—2014-218129

October 2014
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Daniel L. Tweedt

AP Solutions, Inc., Cleveland, Ohio

Computational Aerodynamic Simulations of an

840 ft/sec Tip Speed Advanced Ducted Propulsor 

Fan System Model for Acoustic Methods

Assessment and Development

NASA/CR—2014-218129

October 2014

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared under Contract NNC06BA07B, Task NNC07E190T
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Level of Review: This material has been technically reviewed by NASA technical management OR expert reviewer(s).
Abstract

Computational Aerodynamic simulations of an 840 ft/sec tip speed, Advanced Ducted 

Propulsor fan system were performed at five different operating points on the fan operating line, in 

order to provide detailed internal flow field information for use with fan acoustic prediction 

methods presently being developed, assessed and validated. The fan system is a sub-scale, lownoise research fan/nacelle model that has undergone extensive experimental testing in the 9- by 15-

foot Low Speed Wind Tunnel at the NASA Glenn Research Center, resulting in quality, detailed 

aerodynamic and acoustic measurement data.

Details of the fan geometry, the computational fluid dynamics methods, the computational 

grids, and various computational parameters relevant to the numerical simulations are discussed. 

Flow field results for three of the five operating conditions simulated are presented in order to 

provide a representative look at the computed solutions.

Each of the five fan aerodynamic simulations involved the entire fan system, excluding a

long core duct section downstream of the core inlet guide vane. As a result, only fan rotational

speed and system bypass ratio, set by specifying static pressure downstream of the core inlet guide

vane row, were adjusted in order to set the fan operating point, leading to operating points that lie

on a fan operating line and making mass flow rate a fully dependent parameter. The resulting mass

flow rates are in good agreement with measurement values.

The computed blade row flow fields for all five fan operating points are, in general, 

aerodynamically healthy. Rotor blade and fan exit guide vane flow characteristics are good, 

including incidence and deviation angles, chordwise static pressure distributions, blade surface 

boundary layers, secondary flow structures, and blade wakes. Examination of the computed flow 

fields reveals no excessive boundary layer separations or related secondary-flow problems. A few 

spanwise comparisons between computational and measurement data in the bypass duct show that 

they are in good agreement, thus providing a partial validation of the computational results.

Introduction

The development and validation of aircraft engine fan acoustic prediction methods is an 

important part of ongoing efforts by NASA and industry to reduce noise generation in the fan 

section of aircraft engines. This work is part of a larger task involving computational fluid

dynamics (CFD) to simulate the aerodynamics of selected fan systems, each at several different 

operating points, for the purpose of providing detailed internal flow field information for use with

fan acoustic prediction methods presently being developed, assessed and validated.

This report documents CFD work done on one of the selected fan systems, the Advanced

Ducted Propulsor (ADP) Fan, version 1, which 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. Several years of testing in the LSWT have produced quality, detailed

aerodynamic and acoustic measurement data for the ADP fan [3]á.

Ü Numbers in square brackets indicate references.

NASA/CR—2014-218129 1

Computational Aerodynamic Simulations of an 840 ft/sec Tip Speed 

Advanced Ducted Propulsor Fan System Model for Acoustic 

Methods Assessment and Development 

Daniel L. Tweedt 

AP Solutions, Inc. 

Cleveland, Ohio 44135 
Fan Geometry

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

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

axial location, and the fan exit guide vane (FEGV) row is in the standard/middle configuration 

where its stacking line is 7.400 inches axially downstream of the rotor blade stacking line. The fan 

system also involves a long core-duct section (not shown) downstream of the core inlet guide vane 

(CIGV) row, but it was not included as part of the computational domain. The number of blades 

for each blade row is indicated in parenthesis. The rotating portion of the rotor hub includes the 

entire upstream centerbody/spinner, and extends about one inch downstream of the rotor to axial 

location 2.745 inches (see Figure 1, mark below hub contour downstream of rotor trailing edge).

Rotor blade coordinates for the fan were provided for the running (hot) blade shapes at the 

cruise operating point, which is also the aerodynamic design point. All operating points pertinent 

to this task involve the rotor blades in their sea-level takeoff (SLTO) configuration, which is at a 

blade setting angle closed 9.0 degrees from the cruise angle. At the SLTO rotational speed of 8750 

rpm (mechanical and corrected), corresponding to a rotor blade tip speed of 840 ft/sec, the running 

blade tip deflections are expected to be very nearly the same ñ a difference of 0.033 degrees ñ as 

those for the fan running in the NASA GRC W8 test facility at the cruise operating point (see 

Table VIII in Reference 1), which is the condition at which blade fabrication coordinate 

corrections (hot-to-cold) were made. Therefore, the running blade shape at the SLTO condition in 

the NASA GRC 9- by 15-foot LSWT is, for all practical purposes and based on Reference 1, the 

same as that of the supplied running shape with the blade setting angle reset by a -9.0 degree 

coordinate rotation about the stacking axisß.

Five aerodynamic simulation cases have been defined for the ADP fan, each at a different 

rotational speed on the fan operating line. In all cases mechanical speed is equal to corrected speed 

since the far field flow is at standard day sea-level total (stagnation) conditions. The rotor blade tip 

clearance at the SLTO rotational speed of 8750 rpm is specified to be 0.0200 inches. Approximate 

corrections to rotor blade twist and blade tip clearance were made at the other (lower) rotational 

speeds by assuming blade deflection to be proportional to the square of rotational speed. These 

results are summarized in Table 1. For blade twist corrections, the reported cold-to-hot tip untwist 

value (Reference 1, Table VIII) of 1.166 degrees at 8750 rpm was used to calculate the tip untwist 

values at the other rotational speeds. The spanwise variation in untwist was assumed to be linear, 

varying between zero at the hub and the calculated value at the tip. Tip clearance values were 

estimated based on experimental spin-test results [4] indicating an average tip radius increase of 

0.0145 inches at 10,000 rpm (0.025 inches at the blade leading edge and 0.004 inches at the trailing 

edge).

á The published data were obtained for a configuration with a takeoff-speed rotor tip clearance of 0.055 inches, 

which is over twice as large as the tip clearance for the current task. The experimental measurement data used for 

comparisons in this report are unpublished as of the report submission date.

ß 9.0 degree clockwise rotation when viewed looking radially inward along the stacking axis.

NASA/CR—2014-218129 2
Table 1: Rotor Blade Tip Geometry Corrections

Operating

Point

Rotational

Speed

(rpm)

Blade Tip

Untwist

(degrees)

Tip Untwist

Correction

(degrees)

Blade Tip

Clearance

(inches)

SLTO 8750 1.166 0.000 0.0200

8345 1.061 0.105 0.0210

Cutback 7525 0.862 0.304 0.0229

6700 0.684 0.482 0.0246

Approach 5425 0.448 0.718 0.0268

The rotor hub has two clearance gaps that are separated by the blade trunnion, which 

extends from about 18.5 percent to 72.5 percent of blade chord. These gaps are assumed to be 

invariant with rotational speed. Measurements of the clearance gap heights indicate an average 

leading edge clearance of about 0.0135 inches, and a trailing edge clearance of about 0.0180 

inches. For the CFD simulations the gap heights were approximated using a cubic polynomial for 

the variation of clearance between the leading and trailing edge. The trunnion location differs 

slightly from that of the hardware, mainly because the additional effort to generate grids 

conforming exactly to the hardware was not warranted. Figure 2 provides a graphical summary of 

the hub clearance gap height variation used for the CFD simulations.

Computational Fluid Dynamics

Two different CFD codes were used to simulate the airflow around and through the fan 

system: 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 three-dimensional 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 three-dimensional TWSIFT 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 rowsÜ.

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 the body-normal direction(s), but neglect them in the 

streamwise direction by applying the thin-layer approximation. 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 [5-8]. All simulations described herein were run 

Ü All program-to-program communications, for mixing planes and direct block-to-block interfaces, were handled 

using a facility called SYNCEX (pronounced ìsink-exî). SYNCEX is a message-passing interface that enables 

two or more programs to efficiently exchange data on a single computer and/or over a network.

NASA/CR—2014-218129 3
using a 2-stage Runge-Kutta scheme with a CFL number of 2.5, and using the AUSM+ upwind 

scheme [9] for best accuracy.

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

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

multiblock capability (when used with SYNCEX; see preceding footnote), includes the twoequation SST turbulence model developed by Menter [11], and implements Giles' twodimensional, steady-state, non-reflecting boundary conditions [12,13] at flow inlet, exit, and 

mixing-plane boundariesÜ. Note that when a two-equation turbulence model is used, either the 

Wilcox k- model [14] 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.

Computational Grids

An axisymmetric grid consisting of four two-dimensional grid blocks, shown in Figure 3, 

was used outside of the fan system blade rows. For clarity, only some of the grid lines are drawn, 

and the different blocks are shown in various colors: a far-field block (green), an external nacelle 

block (blue), an upstream/inlet block (black), and a nozzle/downstream block (black). The far-field 

block size is 169×45 nodes, the external nacelle block size is 297×65 nodes, and the upstream/inlet 

and nozzle/downstream block sizes are each 169×85 nodes. The far-field block overlaps the top of 

the nacelle block, but the grid nodes are not aligned, so the computational solutions are 

interpolated there. The bottom of the nacelle block does not overlap the upstream/inlet and 

nozzle/downstream blocks, but the boundary-normal grid spacings are relatively small and the 

boundary nodes are aligned.

The nacelle, inlet, and nozzle grid blocks were all generated using a Poisson partial 

differential equation (PDE) solver, otherwise known as an elliptic grid generator, which produces 

grids with good boundary-normal node clustering and spacings, and generally good local 

orthogonality. Since the CFD method always directly includes the viscous sublayer in the nearwall treatment of turbulent boundary layers ñ wall functions are not used ñ the node spacings at 

solid walls are small. In the inlet and on the external surface of the nacelle the wall-normal spacing 

is nominally 0.0001 inches, whereas in the bypass nozzle the wall-normal spacing is nominally 

between 0.0002 and 0.0003 inches. Corresponding inner-variable wall distances, y

+, are generally 

between 1.0 and 3.0.

An enlarged view of the two-dimensional grid in and around the fan system is shown in 

Figure 4, where every other grid line is drawn. In this figure the elliptic grid stretching for the inlet, 

nozzle, and nacelle grid blocks can be seen more easily. Magnified views of the grid around the 

nacelle leading edge and trailing edge, with every grid line drawn, are shown in Figure 5.

Meridional locations of the three-dimensional blade row grid blocks are shown in Figure 6, 

with flow boundaries indicated by dashed and dash-dotted lines. The blue dash-dotted lines 

indicate grid block direct-interfaces, and the black dashed lines indicate mixing-plane interfaces. 

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

NASA/CR—2014-218129 4
There are four primary grid blocks for the rotor: the rotor inlet H-grid block, the rotor blade row Cgrid block, the rotor exit core H-grid block, and the rotor exit bypass H-grid block. The FEGV and 

CIGV computational domains each involve a single C-grid block. The red streamlines in Figure 6 

show stream-surface locations for blade-to-blade (streamwise-pitchwise) grid views, as well as for 

blade-to-blade flow contour plots, to be shown later.

The blade row grids, except for the two rotor exit H-grid blocks, were generated using a 

computer program called TTGRID, which is a modified version of TCGRID [15], a grid generator 

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

elliptic PDE solver to the blade-to-blade mesh surfaces of blade row C- and H-grids.

Meridional plane projections of the three-dimensional blade row grid blocks at grid 

surfaces located about mid-pitch between the blades are shown in Figure 7, and an enlarged view 

of the grid blocks in the core inlet region is shown in Figure 8. Rotor grid blocks are drawn in 

black and green, and the FEGV and CIGV grid blocks are drawn in red. For clarity, only every 

other grid line in the streamwise direction is shown, although all grid lines in the spanwise 

direction are drawn. Corresponding three-dimensional views of the grid blocks are provided in 

Figures 9 and 10, again with only some of the grid lines drawn. The rotor C-grid (black) has a size 

of 193×49×85 nodes, the rotor inlet H-grid (green) has a size of 21×26×85 nodes, the rotor exit 

core H-grid (green) has a size of 49×89×53 nodes, and the rotor exit bypass H-grid (green) has a 

size of 69×89×73 nodes. Note that the rotor inlet H-grid overlaps the rotor blade C-grid and has 

node-to-node alignment with it. The rotor exit H-grids for the core and bypass also overlap the 

rotor blade C-grid, but the grid nodes are not aligned. On their common boundary just upstream of 

the flow path splitter, the two rotor exit H-grid blocks do not overlap each other, but the boundarynormal spacings are small and the nodes are aligned. The FEGV C-grid block (red) has a size of 

193×45×73 nodes, and the CIGV C-grid block (red) has a size of 177×41×53 nodes. All of the 

three-dimensional grids have boundary-normal node spacings which are nominally 0.0002 inches 

at blade/vane surfaces, and 0.0003 inches at endwall surfaces.

Streamwise-pitchwise views of the rotor grid at three spanwise locations, corresponding to 

the red streamlines in Figure 6, are shown in Figure 11 with every other grid line drawn. 

Corresponding pitchwise-spanwise views of the rotor C-grid and exit H-grid at the respective 

block downstream boundaries are shown in Figure 12.

Over the rotor blade tip, in the endwall clearance gap, an O-grid block of size 161×13×13 

nodes is used, and at the rotor hub two C-grid blocks are used in the clearance gaps, one with a size 

of 55×13×11 nodes for the leading clearance, and the other with a size of 63×13×11 nodes for the 

trailing clearance. The tip clearance grid is shown in Figure 13, which includes magnified views 

around the blade leading- and trailing-edges, and a magnified axial cross-section view near midchord. All grid lines are drawn for the magnified views. The two hub clearance grids are shown in 

Figure 14. Note that each of the three clearance gap grids overlaps the rotor blade C-grid and has 

node-to-node alignment with it.

A streamwise-pitchwise view of the FEGV grid near midspan, corresponding to the middle 

red streamline in Figure 6, is shown in Figure 15. Shown below the full view are magnified views 

of the vane leading and trailing edge regions. For clarity, only every other grid line parallel to the 

vane surface is drawn for the full view. A corresponding pitchwise-spanwise view of the FEGV 

grid at the block downstream boundary is shown in Figure 16. Lastly, a couple of representative 

views of the CIGV grid are shown in Figure 17.

NASA/CR—2014-218129 5
All of the computational grid blocks and their respective sizes are summarized below in 

Tables 2 and 3.

Table 2: Two-Dimensional Grid Blocks

Grid Block Size (I×J×K) Number of Nodes

Fan System Upstream/Inlet 169×85 14,365

Fan System Nozzle/Downstream 169×85 14,365

Fan System External Nacelle 297×65 19,305

Fan System Far Field 169×45 7,605

Total All Blocks 55,640

Table 3: Three-Dimensional Grid Blocks

Grid Block Size (I×J×K) Number of Nodes

Rotor Inlet H-Grid 21×26×85 46,410

Rotor Blade C-Grid 193×49×85 803,845

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

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

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

Rotor Hub Clearance Leading C-Grid 55×13×11 7,865

Rotor Hub Clearance Trailing C-grid 63×13×11 9,009

FEGV C-Grid 193×45×73 634,005

CIGV C-Grid 177×41×53 384,621

Total All Blocks 2,592,390

To conclude this section, the rotor exit bypass H-grid block and its relative importance to 

the CFD solutions will be discussed. It is perhaps apparent that this grid block involves a relatively 

large number of grid nodes, even though it contains no blade surfaces and is essentially just a 

downstream extension of the rotor blade C-grid. The higher grid density is necessary, however, to 

provide the numerical resolution needed for accurate wake convection, and to achieve a reasonable 

level of grid independence for the CFD solution. If the grid is too coarse, particularly in regions 

where the flow field involves large gradients and the primary flow is not aligned with the grid, then 

numerical dissipation is excessive and causes substantial distortion of computed local flow 

features. More specifically, excess artificial dissipation causes the computed wake and blade tip 

vortex to decay too rapidly.

NASA/CR—2014-218129 6
If the primary purpose of the CFD simulations were only aerodynamic performance 

assessment and/or prediction, then the lack of local flow field accuracy in the wake region would 

probably not be crucial because local accuracy typically has a relatively small influence on 

spatially averaged performance quantities. For the current task, however, the computed rotor wake 

and tip-vortex structures are important because they define flow-field characteristics associated 

with noise generation. Particularly important are the computed flow field results at the rotor exit 

bypass H-grid downstream boundary (and mixing plane) since these are intended for direct use in 

acoustic methods assessment, research and development.

Fan System Aerodynamic Simulations

All CFD simulations were run with the far-field (free stream, flight) Mach number set at 

0.100, with total (stagnation) conditions set at standard day sea-level values. The corresponding 

unit Reynolds number is 5.915E+05 inches-1. Air is modeled as a perfect gas with a ratio of 

specific heats, , equal to 1.400.

The effects of turbulence were modeled using the two-equation SST turbulence model [11], 

with free stream turbulence on the far-field upstream boundary set at 0.2 percent, along with a 

turbulence (eddy) viscosity equal to 0.2 times the molecular viscosity, giving a turbulence length 

scale of 1.39E-03 inches. The corresponding turbulence kinetic energy is 5.99E-08 (dimensionless; 

multiply by square of free stream stagnation speed-of-sound to obtain a dimensional value)Ü. In all 

the cases simulated, rotor blade laminar-to-turbulent boundary layer transition occurred near the 

leading edge, at a location around 5 percent of blade chord.

Since the duct downstream of the CIGV row was not modeled in the CFD simulations, it 

was necessary to specify static pressures at the exit boundary of the CIGV grid. 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. At each operating point the value of the hub static 

pressure was determined iteratively so as to achieve the desired system bypass ratio (BPR) for that 

operating point. Measured and computed values of the fan (rotor) mass flow rate and the system 

bypass ratio are listed below in Table 4.

Ü At the fan inlet the turbulence is lower than originally intended because of turbulence decay upstream of the 

inlet, and because of an error in calculating the inlet boundary values so as to account for that decay. The 

turbulence is not so low, however, that it significantly affects the computed aerodynamics. Turbulence at the fan 

inlet is nominally about 0.03 percent, with a turbulence viscosity of 0.2 times the molecular viscosity and a 

turbulence kinetic energy of around 3.5E-08.

NASA/CR—2014-218129 7
Table 4: Measured and Computed Fan Flow Rates and System Bypass Ratios

Operating

Point

Rotational

Speed

(rpm)

Measured

Flow Rate

(lbm/sec)

Computed

Flow Rate

(lbm/sec)

Measured

BPR

Computed

BPR

SLTO 8750 79.20 79.61 11.50 11.49

8345 75.60 76.11 11.20 11.20

Cutback 7525 68.10 68.86 10.70 10.73

6700 60.70 61.55 10.30 10.28

Approach 5425 49.10 50.11 9.90 9.95

All five of the CFD simulation cases were run until the maximum and average (RMS) 

solution residuals were reduced by 4 to 5 orders of magnitude. Similar levels of convergence were 

obtained for performance-related quantities such as mass flow rate, average total temperature, and 

average total pressure. Integrated mass flow conservation discrepancies for the computed solutions 

are small, being everywhere less than ±0.04 percent.

The measured and computed mass flow rates in Table 4 differ at the SLTO condition by 

0.52 percent, and at the approach condition by 2.06 percent. Between those two conditions the 

difference varies almost linearly with rotational speed. The uncertainty in measured flow rate is 

estimated to be at least ±0.5 percent, and could be significantly larger at the lower flow rates 

because the experimental values are calculated from a correlation using static pressures measured 

in the rotor seal cavity. There is, therefore, some uncertainty as to how much the computed and 

measured flow rates actually differ, so the agreement may be somewhat better or worse. In any 

case, the differences are not considered large enough to substantially compromise the intended use 

of the computational results; that is, for acoustic methods assessment, research and development.

Graphs in Figure 18 show spanwise distributions of total pressure and total temperature in 

the bypass duct, upstream of the FEGV, at a location 5.000 inches downstream of the rotor 

stacking line (see gray vertical line in Figure 1). Pitchwise mixed-out-averaged computational 

results are compared with rake measurement data for three of the five fan operating conditions: 

SLTO, cutback, and approach. As can be seen, there is generally good agreement between the 

computational and experimental data.

Selected results from three of the five CFD simulation cases are presented below. Flow 

field contour plots are presented for results at the SLTO, cutback, and approach operating points, 

the primary purpose being to provide a representative look at the computed solutions. More 

extensive and detailed flow field information can be obtained directly from the CFD grid and 

solution data sets, which are being made available along with this report, or which can be obtained 

separately upon request.

Mach number contours for the entire fan system flow field at the SLTO operating point are 

shown in Figure 19, where the three-dimensional blade-row solutions have been mixed-out 

averaged in the pitchwise direction. An enlarged view of the fan region is shown in Figure 20, and 

NASA/CR—2014-218129 8
a corresponding contour plot of turbulence kinetic energy, mass-averaged in the pitchwise 

direction, is shown in Figure 21.

Rotor blade-to-blade relative (rotating system) Mach number contours for the SLTO 

operating point are shown in Figures 22a and 22b. Figure 22a shows a near tip section at about 89 

percent span from the hub, and Figure 22b shows near midspan and near hub sections at about 51 

and 8 percent span, respectively (see red streamlines in Figure 6). Relative Mach number contours 

in and around the rotor tip and hub endwall clearance gaps are shown in Figures 23 and 24, 

respectively, where the tip mid-clearance-gap location is about 0.010 inches from the casing 

endwall, the hub mid-clearance-gap location is about 0.006 inches from the hub endwall, and the 

mid-chord axial location is at the rotor blade stacking line. Corresponding contours of relative 

Mach number at the rotor C-grid downstream boundary, located 2.500 inches downstream of the 

rotor stacking line, are shown in Figure 25, and contours of various flow properties at the rotor exit 

bypass H-grid downstream boundary, which is the downstream mixing plane, are shown in 

Figures 26a and 26b. Figure 26a shows relative and absolute (stationary system) Mach number 

contours, and Figure 26b shows entropy and turbulence kinetic energy contours. Note that the 

entropy is non-dimensionalized by the gas constant, R, and is zero at the upstream reference 

condition.

FEGV blade-to-blade absolute Mach number contours for fan operation at the SLTO 

operating point are shown in Figures 27a and 27b. Figure 27a shows a vane section at about 89 

percent span from the hub, and Figure 27b shows vane sections at about 48 and 10 percent span 

(see red streamlines in Figure 6). Corresponding Mach number contours at the FEGV downstream 

mixing plane are shown in Figure 28.

Mach number contours for the entire fan system at the cutback operating point are shown in 

Figure 29, with an enlarged view of the fan region shown in Figure 30. Again, the threedimensional blade-row solutions have been mixed-out averaged in the pitchwise direction. 

Corresponding rotor blade-to-blade relative Mach number contours are shown in Figures 31a and 

31b, and various flow property contours at the rotor downstream mixing plane in the bypass duct 

are shown in Figures 32a and 32b. Flow field contours for the FEGV at cutback operation are not 

shown, but are aerodynamically similar to those at SLTO.

Mach number contours for the entire fan system at the approach operating point are shown 

in Figure 33, with Figure 34 showing an enlarged view of the fan region. Corresponding rotor 

blade-to-blade relative Mach number contours are shown in Figures 35a and 35b, and various flow 

property contours at the rotor downstream mixing plane in the bypass duct are shown in 

Figures 36a and 36b. Absolute Mach number contours for the FEGV are shown in Figures 37a and 

37b, and in Figure 38. The computed FEGV flow field, like that for cutback operation, is 

aerodynamically similar to the SLTO solution, although it might be noted that the FEGV suctionsurface/hub corner flow separation is significantly weaker for the approach operating point 

(compare Figure37 b to Figure 27b, and Figure 38 to Figure 28). In general, however, despite some 

non-conformity, the similar solutions show that the FEGV flow field scales more-or-less with flow 

rate for all simulated operating points on the fan operating line.

NASA/CR—2014-218129 9
Concluding Remarks

The entire fan system was aerodynamically simulated for five operating points, requiring 

only fan rotational speed and system bypass ratio to be adjusted as the independent parameters 

when setting each operating point. As a result, the computed operating points lie on a fan operating 

line, and mass flow rate is a dependent parameter. Computed and measured fan system mass flow 

rates are in good agreement, indicating indirectly that the computational and experimental fan 

operating lines are nearly the same.

The computed blade row flow fields at all operating points are, in general and as expected, 

aerodynamically healthy. Rotor blade and FEGV flow characteristics are good, including incidence 

and deviation angles, chordwise static pressure distributions (not shown, but can be inferred from 

Mach number distributions), blade surface boundary layers, secondary flow structures, and blade 

wakes. Examination of the computed flow fields reveals no excessive or critical boundary layer 

separations or related secondary flow problems. A few spanwise comparisons between 

computational and measurement data in the bypass duct show that they are in good agreement, thus 

providing a partial validation of the computational results.

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. Mehmed, O., Janetzke, D., ìDeflection Spin Test Results of NASA/P&W Fan 1 Blades,î 

NASA internal memorandum from Structural Dynamics Branch to Chief, Propeller and 

Acoustics Technology Branch, March 20, 1995.

5. Chima, R. V., "Viscous Three-Dimensional Calculations of Transonic Fan Performance," 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).

6. Chima, R. V., and Yokota, J. W., "Numerical Analysis of Three-Dimensional Viscous Flows 

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

7. Tweedt, Daniel L., and Chima, Rodrick V., "Rapid Numerical Simulation of Viscous 

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

8. Tweedt, Daniel L., Chima, Rodrick V., and Turkel, Eli, "Preconditioning for Numerical 

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

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

NASA/CR—2014-218129 10
9. 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).

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

Documentation, Version 300, August 2003.

11. 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).

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

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

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

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

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

Third edition, 2006.

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

Documentation,î Version 300, July, 2003.

NASA/CR—2014-218129 11
Figure 2: Rotor Hub Clearance Gap Height Variation

Figure 1: Schematic Drawing of the ADP 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)

0.0 0.2 0.4 0.6 0.8 1.0

Fraction of Meridional Chord

13

14

15

16

17

18

19

Rotor Hub Clearance Gap Height (mils)

Hardware

CFD Trunnion

(no gap)

CFD

Hardware

NASA/CR—2014-218129 12
Figure 3: Two-Dimensional Axisymmetric CFD Grid Blocks (Entire Computational Domain)

Figure 4: Enlarged Inlet, Nozzle, and Nacelle Grid Blocks

NASA/CR—2014-218129 13
Figure 5: Magnified Nacelle Leading and Trailing Edge Grids

Figure 6: Meridional Location and Extent of Three-Dimensional CFD Grid Blocks

Rotor

 Inlet

H-Grid

Rotor

C-Grid

FEGV

C-Grid

Rotor Exit

Bypass

H-Grid

Rotor Exit

Core

H-Grid

CIGV

C-Grid

NASA/CR—2014-218129 14
Figure 7: Meridional View of Three-Dimensional CFD Grid Blocks (Mid-Pitch Location)

Figure 8: Enlarged Grid Blocks at Core Inlet Region

NASA/CR—2014-218129 15
Figure 9: Three-Dimensional View of Rotor Grid

Figure 10: Three-Dimensional View of FEGV Grid (left) and CIGV Grid (right)

NASA/CR—2014-218129 16
Figure 11: Rotor Grid at Three Spanwise Locations

Rotor Near Mid

Rotor Near Hub

Rotor Near Tip

NASA/CR—2014-218129 17
Figure 12: Rotor Grid at Block Downstream Boundaries

Rotor Exit

H-Grid

Downstream

Boundary

Rotor

C-Grid

Downstream

Boundary

Tip

Hub

Bypass

Core

NASA/CR—2014-218129 18
Figure 13: Rotor Tip Clearance Grid

Trailing Edge

Leading Edge

A ñ A

Blade Tip

Clearance

Gap

A

A

Rotor Tip

NASA/CR—2014-218129 19
Figure 14: Rotor Hub Clearance Grid

Rotor Hub

Hub Trunnion

Leading Clearance

Trailing Clearance

NASA/CR—2014-218129 20
Figure 15: FEGV Grid at Near Midspan Location

FEGV Near Midspan

Leading Edge Trailing Edge

NASA/CR—2014-218129 21
Figure 16: FEGV Grid at Block Downstream Boundary

Tip

Hub

FEGV C-Grid

Downstream

Boundary

NASA/CR—2014-218129 22
Figure 17: CIGV Grid at Near Midspan Location and at Block Downstream Boundary

CIGV Near Midspan

CIGV C-Grid

Downstream

Boundary

Tip

Hub

NASA/CR—2014-218129 23
Figure 18: Computed and Measured Rotor Exit Total-Pressure and Total-Temperature Spanwise 

Distributions in the Bypass Duct , Upstream of FEGV

1.00 1.05 1.10 1.15 1.20 1.25 1.30 1.35

Dimensionless Total Pressure, p t

 / pref

0

20

40

60

80

100

Percent Passage Height

Approach Cutback SLTO

Approach Cutback SLTO

Computed

Measured

1.00 1.02 1.04 1.06 1.08 1.10 1.12

Dimensionless Total Temperature, Tt

 / Tref

0

20

40

60

80

100

Percent Passage Height

NASA/CR—2014-218129 24
Figure 19: Computed Fan System Flow Field for the SLTO Operating Point;

Pitchwise-Averaged Flow; Mach Number Contours

0.0 0.2 0.4 0.6 0.8

Mach Number

SLTO Operating Point

NASA/CR—2014-218129 25
Figure 20: Computed Fan Blade Row Flow Fields for the SLTO Operating Point;

Pitchwise-Averaged Flow; Mach Number Contours

0.0 0.2 0.4 0.6 0.8

Mach Number

NASA/CR—2014-218129 26
Figure 21: Computed Fan Blade Row Flow Fields for the SLTO Operating Point;

Pitchwise-Averaged Flow; Turbulence Kinetic Energy Contours

0.000 0.004 0.008

Turbulence Kinetic Energy

0.002 0.006

NASA/CR—2014-218129 27
Figure 22a: Computed Rotor Flow Field for the SLTO Operating Point;

Blade-to-Blade Relative Mach Number Contours

Rotor Near Tip

(89% Span)

Rotation

Flow

SLTO Operating Point

0.0 0.6 1.0

Relative Mach Number

0.2 0.4 0.8

NASA/CR—2014-218129 28
Figure 22b: Computed Rotor Flow Field for the SLTO Operating Point;

Blade-to-Blade Relative Mach Number Contours

Rotor Near Mid

(51% Span)

Rotor Near Hub

(8% Span)

NASA/CR—2014-218129 29
Figure 23: Computed Rotor Tip Flow Field for the SLTO Operating Point;

Relative Mach Number Contours (Legend in Figure 22a)

Mid-Chord

Axial Location

Mid-Clearance-Gap

Spanwise Location

Rotor Tip Clearance

Rotation

NASA/CR—2014-218129 30
Figure 24: Computed Rotor Hub Flow Field for the SLTO Operating Point;

Relative Mach Number Contours (Legend in Figure 22a)

Mid-Chord

Axial Location

Mid-Clearance-Gap

Spanwise Location

Rotor Hub Clearance

Rotation

NASA/CR—2014-218129 31
Figure 25: Computed Rotor Flow Field for the SLTO Operating Point; Downstream

Boundary of C-Grid; Relative Mach Number Contours

Rotor C-grid

Downstream

Boundary

0.0

0.8

0.2

0.4

0.6

Relative

Mach

Number

Rotation

NASA/CR—2014-218129 32
Figure 26a: Computed Rotor Flow Field for the SLTO Operating Point; Downstream

Mixing Plane; Relative and Absolute Mach Number Contours

Absolute

Mach

Number

0.0

0.2

0.4

0.5

Rotor Exit

Bypass

H-Grid

Downstream

Boundary

Relative

Mach

Number

0.0

0.2

0.4

0.6

Rotation

0.1

0.3

NASA/CR—2014-218129 33
Figure 26b: Computed Rotor Flow Field for the SLTO Operating Point; Downstream

Mixing Plane; Entropy and Turbulence Kinetic Energy Contours

Turbulence

Kinetic

Energy

0.0000

0.0010

0.0015

0.0020

0.0005

Entropy

0.00

0.20

0.25

0.10

Rotor Exit

Bypass

H-Grid

Downstream

Boundary Rotation

0.15

0.05

NASA/CR—2014-218129 34
Figure 27a: Computed FEGV Flow Field for the SLTO Operating Point;

Blade-to-Blade Mach Number Contours

FEGV Near Tip

(89% Span)

Flow

SLTO Operating Point

0.0 0.4 0.6

Mach Number

0.2 0.8

NASA/CR—2014-218129 35
Figure 27b: Computed FEGV Flow Field for the SLTO Operating Point;

Blade-to-Blade Mach Number Contours

FEGV Near Mid

(48% Span)

FEGV Near Hub

(10% Span)

Flow

NASA/CR—2014-218129 36
Figure 28: Computed FEGV Flow Field for the SLTO Operating Point;

Downstream Mixing Plane; Mach Number Contours

FEGV C-Grid

Downstream

Boundary

0.0

0.4

0.1

0.5

Mach

Number

0.2

0.3

NASA/CR—2014-218129 37
Figure 29: Computed Fan System Flow Field for the Cutback Operating Point;

Pitchwise-Averaged Flow; Mach Number Contours

0.0 0.2 0.4 0.6

Mach Number

Cutback Operating Point

NASA/CR—2014-218129 38
Figure 30: Computed Fan Blade Row Flow Fields for the Cutback Operating Point;

Pitchwise-Averaged Flow; Mach Number Contours

0.0 0.2 0.4 0.6

Mach Number

NASA/CR—2014-218129 39
Figure 31a: Computed Rotor Flow Field for the Cutback Operating Point;

Blade-to-Blade Relative Mach Number Contours

Rotor Near Tip

(89% Span)

Rotation

Flow

Cutback Operating Point

0.0 0.6

Relative Mach Number

0.2 0.4 0.8

NASA/CR—2014-218129 40
Figure 31b: Computed Rotor Flow Field for the Cutback Operating Point;

Blade-to-Blade Relative Mach Number Contours

Rotor Near Mid

(51% Span)

Rotor Near Hub

(8% Span)

NASA/CR—2014-218129 41
Figure 32a: Computed Rotor Flow Field for the Cutback Operating Point; Downstream

Mixing Plane; Relative and Absolute Mach Number Contours

Absolute

Mach

Number

0.0

0.2

0.4

Rotor Exit

Bypass

H-Grid

Downstream

Boundary

Relative

Mach

Number

0.0

0.2

0.4

0.6

Rotation

0.5

0.1

0.3

NASA/CR—2014-218129 42
Figure 32b: Computed Rotor Flow Field for the Cutback Operating Point; Downstream

Mixing Plane; Entropy and Turbulence Kinetic Energy Contours

Turbulence

Kinetic

Energy

0.0000

0.0010

0.0015

0.0005

Entropy

0.00

0.10

Rotor Exit

Bypass

H-Grid

Downstream

Boundary Rotation

0.15

0.05

NASA/CR—2014-218129 43
Figure 33: Computed Fan System Flow Field for the Approach Operating Point;

Pitchwise-Averaged Flow; Mach Number Contours

0.0 0.1 0.3 0.4 0.5

Mach Number

Approach Operating Point

0.2

NASA/CR—2014-218129 44
Figure 34: Computed Fan Blade Row Flow Fields for the Approach Operating Point;

Pitchwise-Averaged Flow; Mach Number Contours

0.0 0.1 0.3 0.4 0.5

Mach Number

0.2

NASA/CR—2014-218129 45
Figure 35a: Computed Rotor Flow Field for the Approach Operating Point;

Blade-to-Blade Relative Mach Number Contours

Rotor Near Tip

(89% Span)

Rotation

Flow

Approach Operating Point

0.0 0.6

Relative Mach Number

0.2 0.4

NASA/CR—2014-218129 46
Figure 35b: Computed Rotor Flow Field for the Approach Operating Point;

Blade-to-Blade Relative Mach Number Contours

Rotor Near Mid

(51% Span)

Rotor Near Hub

(8% Span)

NASA/CR—2014-218129 47
Figure 36a: Computed Rotor Flow Field for the Approach Operating Point; Downstream

Mixing Plane; Relative and Absolute Mach Number Contours

Absolute

Mach

Number

0.0

0.2

Rotor Exit

Bypass

H-Grid

Downstream

Boundary

Relative

Mach

Number

0.0

0.2

0.4

Rotation

0.5

0.1

0.3

0.3

0.1

NASA/CR—2014-218129 48
Figure 36b: Computed Rotor Flow Field for the Approach Operating Point; Downstream

Mixing Plane; Entropy and Turbulence Kinetic Energy Contours

Turbulence

Kinetic

Energy

0.0000

0.0006

0.0004

Entropy

0.00

0.06

Rotor Exit

Bypass

H-Grid

Downstream

Boundary Rotation

0.08

0.02

0.04

0.0002

NASA/CR—2014-218129 49
Figure 37a: Computed FEGV Flow Field for the Approach Operating Point;

Blade-to-Blade Mach Number Contours

FEGV Near Tip

(89% Span)

Flow

Approach Operating Point

0.0 0.3 0.4

Mach Number

0.1 0.2 0.5

NASA/CR—2014-218129 50
Figure 37b: Computed FEGV Flow Field for the Approach Operating Point;

Blade-to-Blade Mach Number Contours

FEGV Near Mid

(48% Span)

FEGV Near Hub

(10% Span)

Flow

NASA/CR—2014-218129 51
Figure 38: Computed FEGV Flow Field for the Approach Operating Point;

Downstream Mixing Plane; Mach Number Contours

FEGV C-Grid

Downstream

Boundary

0.0

0.1

Mach

Number

0.2

0.3

NASA/CR—2014-218129 52

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4. TITLE AND SUBTITLE

Computational Aerodynamic Simulations of an 840 ft/sec Tip Speed Advanced Ducted

Propulsor Fan System Model for Acoustic Methods Assessment and Development

5a. CONTRACT NUMBER

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Tweedt, Daniel, L.

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WBS 473452.02.03.07.06.01.06

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Subject Category: 71

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

This publication is available from the NASA Center for AeroSpace Information, 443-757-5802

13. SUPPLEMENTARY NOTES

14. ABSTRACT

Computational Aerodynamic simulations of an 840 ft/sec tip speed, Advanced Ducted Propulsor fan system were performed at five different operating points on the fan

operating line, in order to provide detailed internal flow field information for use with fan acoustic prediction methods presently being developed, assessed and validated. The

fan system is a sub-scale, lownoise research fan/nacelle model that has undergone extensive experimental testing in the 9- by 15-foot Low Speed Wind Tunnel at the NASA

Glenn Research Center, resulting in quality, detailed aerodynamic and acoustic measurement data. Details of the fan geometry, the computational fluid dynamics methods, the

computational grids, and various computational parameters relevant to the numerical simulations are discussed. Flow field results for three of the five operating conditions

simulated are presented in order to provide a representative look at the computed solutions. Each of the five fan aerodynamic simulations involved the entire fan system,

excluding a long core duct section downstream of the core inlet guide vane. As a result, only fan rotational speed and system bypass ratio, set by specifying static pressure

downstream of the core inlet guide vane row, were adjusted in order to set the fan operating point, leading to operating points that lie on a fan operating line and making mass

flow rate a fully dependent parameter. The resulting mass flow rates are in good agreement with measurement values. The computed blade row flow fields for all five fan

operating points are, in general, aerodynamically healthy. Rotor blade and fan exit guide vane flow characteristics are good, including incidence and deviation angles,

chordwise static pressure distributions, blade surface boundary layers, secondary flow structures, and blade wakes. Examination of the computed flow fields reveals no

excessive boundary layer separations or related secondary-flow problems. A few spanwise comparisons between computational and measurement data in the bypass duct show

that they are in good agreement, thus providing a partial validation of the computational results.

15. SUBJECT TERMS

Low noise; Models; Scale models; Turbofans; Acoustic measurement; Aerodynamic

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