Daniel L. Tweedt

AP Solutions, Inc., Cleveland, Ohio

Computational Aerodynamic Simulations of a

1215 ft/sec Tip Speed Transonic Fan System Model 

for Acoustic Methods Assessment and Development

NASA/CR—2014-218130

October 2014
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Daniel L. Tweedt

AP Solutions, Inc., Cleveland, Ohio

Computational Aerodynamic Simulations of a

1215 ft/sec Tip Speed Transonic Fan System Model 

for Acoustic Methods Assessment and Development

NASA/CR—2014-218130

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 a 1215 ft/sec tip speed transonic 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, low-noise 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.

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 points 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, which for this 

model did not include a split flow path with core and bypass ducts. As a result, it was only

necessary to adjust fan rotational speed 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.

Computed blade row flow fields at all 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 flow fields at all operating conditions reveals

no excessive boundary layer separations or related secondary-flow problems.

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 Source 

Diagnostics Test (SDT) fan with rotor R4 and a standard, short-chord stator, both of which were 

designed by the General Electric Corporation with partial funding from NASA under the Advanced 

Subsonic Technology program. The SDT 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 [1,2]Ü.

Ü Numbers in square brackets indicate references.

NASA/CR—2014-218130 1

Computational Aerodynamic Simulations of a 1215 ft/sec 

Tip Speed Transonic 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 SDT 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 a short-chord, high vane-count, standard 

design with the trailing-edge stacking line located 8.807 inches axially downstream of the rotor 

blade stacking line. 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 0.150 inches downstream of the rotor to axial location 1.640 inches (see Figure 1, mark 

below hub contour near rotor trailing edge).

Five aerodynamic simulation cases were defined for the SDT 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. Rotor blade 

coordinates were provided for the running (hot) blade shape at three of the five fan operating 

points: sea-level takeoff (SLTO), cutback, and approach, corresponding to corrected rotational 

speeds of 12657, 11075, and 7809 rpm, respectively. Blade coordinates for the other two rotational 

speeds, 11771 and 9493 rpm, were determined by linear interpolation. The corrected rotor blade tip 

speed for the SLTO operating point is 1215 ft/sec. Running rotor tip clearances were measured at 

blade leading edge, mid-chord, and trailing edge locations, over the entire range of rotational 

speeds, and quadratic regression curve fits of the measured data were used to determine the 

clearances at each of the selected rotational speeds. Table 1 summarizes these results. Quadratic 

functions defined from the data in Table 1 were used for calculating the chordwise distributions of 

tip clearance, which are shown in Figure 2.

Table 1: Rotor Blade Tip Clearances

Operating

Point

Rotational

Speed

(rpm)

Leading Edge

Tip Clearance

(mils)

Mid-Chord

Tip Clearance

(mils)

Trailing Edge

Tip Clearance

(mils)

SLTO 12,657 34.87 31.86 39.74

11,771 36.40 34.08 41.82

Cutback 11,075 37.57 35.75 43.42

9,493 40.09 39.35 46.96

Approach 7,809 42.58 42.86 50.60

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 

NASA/CR—2014-218130 2
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 [3-6]. All simulations described herein were run 

using a 2-stage Runge-Kutta scheme with a CFL number of 2.5, and using the AUSM+ upwind 

scheme [7] for best accuracy.

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

[8] 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 [9], and implements Giles' two-dimensional, 

steady-state, non-reflecting boundary conditions [10,11] at flow inlet, exit, and mixing-plane 

boundariesá. Note that when a two-equation turbulence model is used, either the Wilcox k- model 

[12] 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 massaveraged, 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 177×45 nodes, the external nacelle block size is 321×65 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 nearÜ 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 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—2014-218130 3
wall 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 nozzle the wall-normal spacing is nominally 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 

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

the nacelle leading edge and trailing edge 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. 

There are three primary grid blocks for the rotor: the rotor inlet H-grid block, the rotor blade row 

C-grid block, and the rotor exit H-grid block. The FEGV computational domain involves 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 rotor exit H-grid block, were generated using a 

computer program called TTGRID, which is a modified version of TCGRID [13], 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 Figures 7. Rotor grid blocks are 

drawn in black and green, and the FEGV grid is 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 8 and 9, 

again with only some of the grid lines drawn. The rotor C-grid (black) has a size of 217×49×85 

nodes, the rotor inlet H-grid (green) has a size of 33×42×85 nodes, and the rotor exit H-grid 

(green) has a size of 153×89×81 nodes. Note that the rotor inlet H-grid overlaps the rotor blade Cgrid and has node-to-node alignment with it. The rotor exit H-grid also overlaps the rotor blade Cgrid, but the grid nodes are not aligned. The FEGV C-grid (red) has a size of 209×45×73 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 10, with every other grid line drawn (every 

fourth line in the streamwise direction for the exit H-grid block). Corresponding pitchwisespanwise views of the rotor C-grid and exit H-grid at the respective block downstream boundaries 

are shown in Figure 11.

Over the rotor blade tip, in the endwall clearance gap, an O-grid block of size 179×13×17 

nodes is used. The tip clearance grid is shown in Figure 12, 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. Note that the tip clearance grid overlaps 

the rotor blade C-grid and has node-to-node alignment with it.

NASA/CR—2014-218130 4
A streamwise-pitchwise view of the FEGV grid near midspan, corresponding to the middle 

red streamline in Figure 6, is shown in Figure 13. 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 14.

All of all 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 321×65 20,865

Fan System Far Field 177×45 7,965

Total All Blocks 57,560

Table 3: Three-Dimensional Grid Blocks

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

Rotor Inlet H-Grid 33×42×85 117,810

Rotor Blade C-Grid 217×49×85 903,805

Rotor Exit H-Grid 153×89×81 1,102,977

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

FEGV C-Grid 209×45×73 686,565

Total All Blocks 2,850,716

To conclude this section, the rotor exit 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—2014-218130 5
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 might 

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 tipvortex 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 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 [9], 

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.

Measured and computed values of the fan mass flow rate are listed below in Table 4. 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 ±0.06 percent.

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—2014-218130 6
Table 4: Measured and Computed Fan Flow Rates

Operating

Point

Rotational

Speed

(rpm)

Measured

Flow Rate

(lbm/sec)

Computed

Flow Rate

(lbm/sec)

Flow Rate

Difference

(percent)

SLTO 12,657 97.34 97.23 -0.11

11,771 89.66 89.53 -0.14

Cutback 11,075 83.89 83.48 -0.49

9,493 70.83 70.52 -0.44

Approach 7,809 57.86 57.97 0.19

The measured and computed mass flow rates in Table 4 differ by -0.49 to 0.19 percent, 

depending on operating point. The uncertainty in measured flow rate is estimated to be at least 

±0.5 percent, so there is some uncertainty regarding exactly how much the computed and measured 

flow rates differ. In any case, the differences are not large enough to substantially compromise the 

intended use of the computational results; that is, for acoustic methods assessment, research and 

development.

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 15, 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 16, and 

a corresponding contour plot of turbulence kinetic energy, mass-averaged in the pitchwise 

direction, is shown in Figure 17.

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

operating point are shown in Figures 18a and 18b. Figure 18a shows a near tip section at about 89 

percent span from the hub, and Figure 18b shows near midspan and near hub sections at about 51 

and 11 percent span, respectively (see red streamlines in Figure 6). Relative Mach number 

contours in and around the rotor tip endwall clearance gap are shown in Figure 19, where the tip 

mid-clearance-gap location is roughly 0.020 inches from the casing 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 0.550 inches downstream of the rotor trailing 

edge, are shown in Figure 20, and contours of various flow properties at the rotor exit H-grid 

downstream boundary, which is the downstream mixing plane, are shown in Figures 21a and 21b. 

Figure 21a shows relative and absolute (stationary system) Mach number contours, and Figure 21b 

shows entropy and turbulence kinetic energy contours. Note that the entropy is nondimensionalized by the gas constant, R, and is zero at the upstream reference condition.

NASA/CR—2014-218130 7
FEGV blade-to-blade absolute Mach number contours for fan operation at the SLTO 

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

percent span from the hub, and Figure 22b shows vane sections at about 51 and 11 percent span 

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

mixing plane are shown in Figure 23.

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

Figure 24, with an enlarged view of the fan region shown in Figure 25. 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 26a and 

26b, and various flow property contours at the rotor downstream mixing plane are shown in 

Figures 27a and 27b. 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 28, with Figure 29 showing an enlarged view of the fan region. Corresponding rotor 

blade-to-blade relative Mach number contours are shown in Figures 30a and 30b, and various flow 

property contours at the rotor downstream mixing plane are shown in Figures 31a and 31b. 

Absolute Mach number contours for the FEGV are shown in Figures 32a and 32b, and in 

Figure 33. 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 suction-surface/hub corner 

flow separation is significantly weaker for the approach operating point (compare Figure 32b to 

Figure 22b, and Figure 33 to Figure 23). 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.

Concluding Remarks

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

only fan rotational speed to be adjusted as the independent parameter 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.

NASA/CR—2014-218130 8
References

1. 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.

2. 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.

3. 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).

4. 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.

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

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

6. 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.

7. 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).

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

Documentation, Version 300, August 2003.

9. 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).

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

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

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

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

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

Third edition, 2006.

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

Documentation,î Version 300, July, 2003.

NASA/CR—2014-218130 9
Figure 2: Chordwise Distributions of Rotor Blade Tip Clearance

Figure 1: Schematic Drawing of the SDT Fan System

-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)

0 20 40 60 80 100

Meridional Chordwise Location (percent)

30

40

50

Rotor Tip Clearance Gap Height (mils)

12,657

7,809

9,493

11,075

11,771

Rotational

 Speed

 (rpm)

NASA/CR—2014-218130 10
Figure 3: Two-Dimensional Axisymmetric CFD Grid Blocks (Entire Computational Domain)

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

NASA/CR—2014-218130 11
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 Rotor Exit C-Grid H-Grid

NASA/CR—2014-218130 12
Figure 7: Meridional View of Three-Dimensional CFD Grid Blocks (Mid-Pitch Location)

NASA/CR—2014-218130 13
Figure 8: Three-Dimensional View of Rotor Grid

Figure 9: Three-Dimensional View of FEGV Grid

NASA/CR—2014-218130 14
Figure 10: Rotor Grid at Three Spanwise Locations

Rotor Near Mid

Rotor Near Hub

Rotor Near Tip

NASA/CR—2014-218130 15
Figure 11: Rotor Grid at Block Downstream Boundaries

Rotor Exit

H-Grid

Downstream

Boundary

Rotor

C-Grid

Downstream

Boundary

Tip

Hub

NASA/CR—2014-218130 16
Figure 12: Rotor Tip Clearance Grid

Trailing Edge

Leading Edge

A ñ A

Blade Tip

Clearance

Gap

A

A

Rotor Tip

NASA/CR—2014-218130 17
Figure 13: FEGV Grid at Near Midspan Location

FEGV Near Midspan

Leading Edge Trailing Edge

NASA/CR—2014-218130 18
Figure 14: FEGV Grid at Block Downstream Boundary

Tip

Hub

FEGV C-Grid

Downstream

Boundary

NASA/CR—2014-218130 19
Figure 15: 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

1.0

NASA/CR—2014-218130 20
Figure 16: 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

1.0

NASA/CR—2014-218130 21
Figure 17: Computed Fan Blade Row Flow Fields for the SLTO Operating Point;

Pitchwise-Averaged Flow; Turbulence Kinetic Energy Contours

0.000 0.005 0.010

Turbulence Kinetic Energy

NASA/CR—2014-218130 22
Figure 18a: 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 0.5 1.0 1.5

Relative Mach Number

NASA/CR—2014-218130 23
Figure 18b: Computed Rotor Flow Field for the SLTO Operating Point;

Blade-to-Blade Relative Mach Number Contours

Rotor Near Mid

(51% Span)

Rotor Near Hub

(11% Span)

NASA/CR—2014-218130 24
Figure 19: Computed Rotor Tip Flow Field for the SLTO Operating Point;

Relative Mach Number Contours (Legend in Figure 18a)

Mid-Chord

Axial Location

Mid-Clearance-Gap

Spanwise Location

Rotor Tip Clearance

Rotation

NASA/CR—2014-218130 25
Figure 20: 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

1.0

NASA/CR—2014-218130 26
Figure 21a: 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

0.6

Rotor Exit

H-Grid

Downstream

Boundary

Relative

Mach

Number

Rotation

0.0

0.8

0.2

0.4

0.6

1.0

NASA/CR—2014-218130 27
Figure 21b: Computed Rotor Flow Field for the SLTO Operating Point; Downstream

Mixing Plane; Entropy and Turbulence Kinetic Energy Contours

Turbulence

Kinetic

Energy

0.0000

0.0025

0.0010

0.0015

0.0020

0.0005

Entropy

0.0

0.2

0.3

0.1 Rotor Exit

H-Grid

Downstream

Boundary

Rotation

NASA/CR—2014-218130 28
Figure 22a: 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 0.8

Mach Number

0.2 1.0

NASA/CR—2014-218130 29
Figure 22b: Computed FEGV Flow Field for the SLTO Operating Point;

Blade-to-Blade Mach Number Contours

FEGV Near Mid

(51% Span)

FEGV Near Hub

(11% Span)

Flow

NASA/CR—2014-218130 30
Figure 23: 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

0.6

NASA/CR—2014-218130 31
Figure 24: Computed Fan System Flow Field for the Cutback Operating Point;

Pitchwise-Averaged Flow; Mach Number Contours

Cutback Operating Point

0.0 0.2 0.4 0.6 0.8

Mach Number

NASA/CR—2014-218130 32
Figure 25: 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 0.8

Mach Number

NASA/CR—2014-218130 33
Figure 26a: 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 0.5 1.0 1.5

Relative Mach Number

NASA/CR—2014-218130 34
Figure 26b: Computed Rotor Flow Field for the Cutback Operating Point;

Blade-to-Blade Relative Mach Number Contours

Rotor Near Mid

(51% Span)

Rotor Near Hub

(11% Span)

NASA/CR—2014-218130 35
Figure 27a: 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

0.6

Rotor Exit

H-Grid

Downstream

Boundary

Relative

Mach

Number

Rotation

0.0

0.8

0.2

0.4

0.6

NASA/CR—2014-218130 36
Figure 27b: 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

0.0020

0.0005

Entropy

0.0

0.2

0.3

0.1 Rotor Exit

H-Grid

Downstream

Boundary

Rotation

NASA/CR—2014-218130 37
Figure 28: Computed Fan System Flow Field for the Approach Operating Point;

Pitchwise-Averaged Flow; Mach Number Contours

Approach Operating Point

0.0 0.2 0.4 0.5

Mach Number

0.1 0.3

NASA/CR—2014-218130 38
Figure 29: Computed Fan Blade Row Flow Fields for the Approach Operating Point;

Pitchwise-Averaged Flow; Mach Number Contours

0.0 0.2 0.4 0.5

Mach Number

0.1 0.3

NASA/CR—2014-218130 39
Figure 30a: 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 0.8 1.0

Relative Mach Number

0.2 0.4

NASA/CR—2014-218130 40
Figure 30b: Computed Rotor Flow Field for the Approach Operating Point;

Blade-to-Blade Relative Mach Number Contours

Rotor Near Mid

(51% Span)

Rotor Near Hub

(11% Span)

NASA/CR—2014-218130 41
Figure 31a: 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

0.3

0.4

Rotor Exit

H-Grid

Downstream

Boundary

Relative

Mach

Number

Rotation

0.0

0.2

0.4

0.6

0.1

NASA/CR—2014-218130 42
Figure 31b: 

Figure 31b: Computed Rotor Flow Field for the Approach 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

0.20

0.15

0.10

Rotor Exit

H-Grid

Downstream

Boundary

Rotation

0.05

NASA/CR—2014-218130 43
Figure 32a: 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 0.5

Mach Number

0.1 0.2

NASA/CR—2014-218130 44
Figure 32b: Computed FEGV Flow Field for the Approach Operating Point;

Blade-to-Blade Mach Number Contours

FEGV Near Mid

(51% Span)

FEGV Near Hub

(11% Span)

Flow

NASA/CR—2014-218130 45
Figure 33: 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

0.4

NASA/CR—2014-218130 46

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4. TITLE AND SUBTITLE

Computational Aerodynamic Simulations of a 1215 ft/sec Tip Speed Transonic Fan

System Model for Acoustic Methods Assessment and Development

5a. CONTRACT NUMBER

NNC06BA07B

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6. AUTHOR(S)

Tweedt, Daniel, L.

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WBS 473452.02.03.07.06.01.06

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

13. SUPPLEMENTARY NOTES

14. ABSTRACT

Computational Aerodynamic simulations of a 1215 ft/sec tip speed transonic 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, low-noise 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. 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 points 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, which for this model did not include a

split flow path with core and bypass ducts. As a result, it was only necessary to adjust fan rotational speed 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. Computed blade row flow fields at all 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 flow

fields at all operating conditions reveals no excessive boundary layer separations or related secondary-flow problems.

15. SUBJECT TERMS

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

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