Daniel L. Tweedt

AP Solutions, Inc., Cleveland, Ohio

Computational Aerodynamic Simulations of a

1484 ft/sec Tip Speed Quiet High-Speed Fan

System Model for Acoustic Methods Assessment 

and Development

NASA/CR—2014-218131

October 2014
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Daniel L. Tweedt

AP Solutions, Inc., Cleveland, Ohio

Computational Aerodynamic Simulations of a

1484 ft/sec Tip Speed Quiet High-Speed Fan

System Model for Acoustic Methods Assessment 

and Development

NASA/CR—2014-218131

October 2014

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared under Contract NNC06BA07B, Task NNC07E190T
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Trade names and trademarks are used in this report for identifi cation 

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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 1484 ft/sec tip speed quiet high-speed 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 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 

includes a core duct and a bypass duct that merge upstream of the fan system nozzle. As a result, 

only fan rotational speed and the system bypass ratio, set by means of a translating nozzle plug, 

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.

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 computed flow fields reveals no excessive or 

critical boundary layer separations or related secondary-flow problems, with the exception of the 

hub boundary layer at the core duct entrance. At that location a significant flow separation is 

present. The region of local flow recirculation extends through a mixing plane, however, which for 

the particular mixing-plane model used is now known to exaggerate the recirculation. In any case, 

the flow separation has relatively little impact on the computed rotor and FEGV flow fields.

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, version II of the 

Quiet High-Speed (QHS) fan designed by Honeywell Corporation [1]Ü with partial funding from 

NASA under the Advanced Subsonic Technology program. The QHS fan is a 22-inch sub-scale, 

low-noise research fan/nacelle model that has undergone experimental testing in the 9- by 15-foot 

Low Speed Wind Tunnel (LSWT) at the NASA Glenn Research Center.

Ü Numbers in square brackets indicate references.

NASA/CR—2014-218131 1

Computational Aerodynamic Simulations of a 1484 ft/sec Tip Speed 

Quiet High-Speed 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 QHS fan system is shown in Figures 1a and 1b, with all 

major components depicted and shown to scale. The number of blades for each blade row is 

indicated in parenthesis, and the fan rotor blade stacking line is the zero-reference axial location. 

Downstream of the fan exit guide vane (FEGV) row the flow path splits to form a core duct and a 

bypass duct, which merge further downstream ahead of the exit nozzle. Support struts are located 

in the core and bypass ducts, near the flow path split. 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 1.825 inches (see Figure 1b, mark below hub contour near rotor trailing edge).

The fan system bypass ratio (BPR) is controlled by the axial position of the nozzle 

centerbody, or so-called plug, which translates axially to adjust the flow area near the exit of the 

core duct, but without changing the flow area at the nozzle exit. Table 1 lists experimental and 

computational bypass ratios for corresponding nozzle plug positions. Note that a plug position of 

zero corresponds to a fully closed/blocked core duct, and that the computational BPR values were 

obtained by iteratively adjusting the plug position in order to achieve conditions close to the 

corresponding experimental BPR values.

Experimental testing of the QHS fan involved, in general, some different nozzle exit areas, 

each for essentially the same nozzle, but with the nozzle/nacelle trailing edge ìtrimmedî by 

different amounts. More trim provides a larger nozzle exit area, and different exit areas produce 

different fan operating lines. All experimental work relevant to the current task was done using a 

single trim configuration, denoted here as trim-1. To better match the experimental mass flow 

rates, however, a single non-experimental trim configuration, denoted as trim-A, was used for the 

computational simulations. In Figure 2 the geometry of the trim-A configuration is shown and 

compared to the trim-1 configuration. An untrimmed configuration, trim-0, is also shown for 

reference. Note that the nozzle exit area for trim-A is 0.26 percent less than that for trim-1.

Table 1: Nozzle Plug Positions and Corresponding Fan System Bypass Ratios

Operating

Point

Rotational

Speed

(rpm)

Experimental

Plug-Position

(inches)

Computational

Plug-Position

(inches)

Experimental

BPR

Computational

BPR

SLTO 14,060 5.421 4.771 4.454 4.436

13,280 4.931 4.541 4.581 4.615

Cutback 12,500 4.559 4.415 4.737 4.791

10,935 4.180 4.171 4.977 4.965

Approach 9,840 3.959 4.083 5.100 5.085

Five aerodynamic simulation cases were defined for the QHS fan (see Table 1), 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 the design rotational 

NASA/CR—2014-218131 2
speed, and no blade shape corrections were made for other rotational speeds. Note that the 

aerodynamic design point, with a rotational speed of 15621 rpm and a corresponding rotor blade 

tip speed of 1484 ft/sec, was not simulated.

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 2 summarizes these results. Quadratic functions defined from the data in Table 2 were used 

for calculating chordwise distributions of the tip clearance, which are shown in Figure 3.

Table 2: 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 14,060 33.92 26.00 26.70

13,280 35.91 28.77 30.30

Cutback 12,500 37.74 31.43 33.86

10,935 40.96 36.40 40.84

Approach 9,840 42.85 39.60 45.60

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 

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 [2-5]. 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 executing programs to efficiently exchange data on a single computer and/or over a network.

NASA/CR—2014-218131 3
using a 2-stage Runge-Kutta scheme with a CFL number of 2.5, and using the AUSM+ upwind 

scheme [6] for best accuracy.

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

[7] 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 [8], and implements Giles' two-dimensional, 

steady-state, non-reflecting boundary conditions [9,10] at flow inlet, exit, and mixing-plane 

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

[11] 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 six two-dimensional grid blocks, shown in Figure 4, 

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

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

external nacelle block (blue), an upstream/inlet block (black), a nozzle/downstream block (black), 

a bypass duct block (red), and a core duct block (magenta). The far-field block size is 193×45 

nodes, the external nacelle block size is 361×65 nodes, the upstream/inlet block size is 185×85 

nodes, the nozzle/downstream block size is 153×85 nodes, the bypass duct block size is 125×65 

nodes, and the core duct block size is 137×49 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 bypass and core duct blocks overlap the nozzle block, without nodal alignment, while 

along their common boundary they have nodal alignment but do not overlap each other.

The nacelle, inlet, duct, 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 ducts and 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.

Enlarged views of the two-dimensional grid in and around the fan system are shown in 

Figures 5a, 5b, and 5c, where every other grid line is drawn. In these figures the elliptic grid 

stretching for the nacelle, inlet, duct, and nozzle 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 6.

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

NASA/CR—2014-218131 4
Meridional locations of the three-dimensional blade row grid blocks are shown in Figure 7, 

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 also involves three 

grid blocks: the FEGV C-grid block, the FEGV exit core H-grid block, and the FEGV exit bypass 

H-grid block. The red streamlines in Figure 7 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 and FEGV exit H-grid blocks, were generated 

using a computer program called TTGRID, which is a modified version of TCGRID [12], 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 8. Rotor grid blocks are 

drawn in black and green, the FEGV grid blocks are drawn in red and blue, and the core and 

bypass strut grid blocks are drawn in magenta. For clarity, only every other grid line in the 

streamwise direction (every fourth line for the rotor exit block) 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 for the rotor and stator, respectively, again with only some of the grid 

lines drawn. The rotor C-grid (black) has a size of 185×49×85 nodes, the rotor inlet H-grid (green) 

has a size of 25×28×85 nodes, and the rotor exit H-grid (green) has a size of 97×89×81 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-grid also overlaps the rotor blade C-grid, but the grid nodes are not 

aligned. For the FEGV, the C-grid (red) has a size of 193×45×85 nodes, the exit core H-grid (blue) 

has a size of 47×89×57 nodes, and the exit bypass H-grid (blue) has a size of 47×89×73 nodes. As 

with the rotor, the FEGV exit H-grids overlap the C-grid, but the grid nodes are not aligned. On 

their common boundary just upstream of the flow path splitter, the two FEGV exit H-grid blocks 

do not overlap each other, but the boundary-normal spacings are small and the nodes are aligned. 

Lastly, the core strut H-grid has a size of 97×73×49 nodes, and the bypass strut H-grid has a size 

of 97×81×65 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 7, are shown in Figure 11 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 12.

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 13, where magnified views around the 

blade leading- and trailing-edges are included, as well as a magnified axial cross-section view near 

mid-chord. 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.

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

red streamline in Figure 7, is shown in Figure 14. Shown below the full view are magnified views 

of the vane leading and trailing edge regions. In the full view, for clarity, only every other grid line 

NASA/CR—2014-218131 5
parallel to the vane surface is drawn for the C-grid block (red), and every other line in both grid 

directions is drawn for the exit H-grid block (blue). Corresponding pitchwise-spanwise views of 

the FEGV C-grid and exit H-grid at the block downstream boundaries are shown in Figure 15.

A streamwise-pitchwise view of the core-strut grid near midspan, with magnified views of 

the strut leading and trailing edge regions, is shown in Figure 16. Midspan in this case corresponds 

to the lower red streamline in Figure 7. Because the bypass-strut H-grid is similar to the core-strut 

H-grid, only one of them is shown for the blade-to-blade view. Views of both grids at the block 

downstream boundaries, however, are shown in Figure 17.

All computational grid blocks and their respective sizes are summarized below in Tables 3 

and 4.

Table 3: Two-Dimensional Grid Blocks

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

Fan System Upstream/Inlet 185×85 15,725

Fan System Nozzle/Downstream 153×85 13,005

Fan System Core Duct 137×49 6,713

Fan System Bypass Duct 125×65 8,125

Fan System External Nacelle 361×65 23,465

Fan System Far Field 193×45 8,685

Total All Blocks 75,718

Table 4: Three-Dimensional Grid Blocks

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

Rotor Inlet H-Grid 25×28×85 59,500

Rotor Blade C-Grid 185×49×85 770,525

Rotor Exit H-Grid 97×89×81 699,273

Rotor Tip Clearance O-Grid 149×13×17 32,929

FEGV C-Grid 193×45×85 738,225

FEGV Exit Core H-grid 47×89×57 238,431

FEGV Exit Bypass H-grid 47×89×73 305,359

Core Strut H-grid 97×73×49 346,969

Bypass Strut H-grid 97×81×65 510,705

Total All Blocks 3,701,916

NASA/CR—2014-218131 6
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.

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 exit 

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 [8], 

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 5. 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 

fairly small, being normally less than ±0.07 percent at all mixing planes. At the two lowest 

rotational speeds, however, the discrepancies for mixing planes downstream of the splitter are a 

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—2014-218131 7
little larger: around -0.08 percent for the 10935 rpm operating point, and -0.13 percent for 9840 

rpm operating point.

Table 5: 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 14,060 89.64 89.71 0.08

13,280 84.03 84.12 0.11

Cutback 12,500 78.79 78.58 -0.27

10,935 68.21 68.07 -0.21

Approach 9,840 61.00 60.80 -0.33

The measured and computed mass flow rates in Table 5 differ by -0.33 to 0.11 percent, 

depending on operating point. The uncertainty in measured flow rate is estimated to be around ±0.5 

percent. Recall that the nozzle exit area was adjusted by changing the amount of nacelle trailingedge trim, to better match the computed and experimental flow rates. The adjustment, determined 

only at the SLTO operating point, is relatively small, however, with the computational model 

having only 0.26 percent less nozzle exit area than the experimental model.

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 18, 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 19, and 

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

direction, is shown in Figure 20.

Notice in Figure 19 that the hub boundary layer for the (pitchwise-averaged) flow at the 

core duct entrance is separated, with the region of reversed flow extending from just upstream of 

the core entrance mixing plane to about the leading edge of the core strut. The separation starts in 

the FEGV computational domain, is highly three-dimensional, and extends in the circumferential 

direction over only part of the vane pitch. The occurrence of recirculating flow at the mixing plane, 

with flow traversing the plane both ways at particular spanwise locations, presented a significant 

challenge for the mixing-plane computational model. Although the difficulty has since been 

resolved, at the time of the simulations it was handled using only a marginally acceptable scheme. 

Consequently, as is now known, the computed flow recirculation is artificially excessive, or 

exaggerated. Its impact, however, on the overall simulated solutions is relatively minor, in large 

part because the flow separation has little effect on the fan rotor and stator flow fields.. Even at the 

NASA/CR—2014-218131 8
lower fan rotational speeds and flow rates, where the computed recirculation zone is markedly 

larger, the computed fan solutions are considered to be only minimally compromised. Perhaps of 

greater significance is the observation that system bypass ratio was fairly sensitive to the severity 

of the flow separation, making it difficult to closely match (by adjusting the nozzle plug position) 

computational values to experimental ones. This may explain, at least in part, the different 

experimental and computational plug positions at each operating point (see Table 1).

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

operating point are shown in Figures 21a and 21b. Figure 21a shows a near tip section at about 87 

percent span from the hub, and Figure 21b shows near midspan and near hub sections at about 52 

and 15 percent span, respectively (see red streamlines in Figure 7). Relative Mach number 

contours in and around the rotor tip endwall clearance gap are shown in Figure 22, where the tip 

mid-clearance-gap location is roughly 0.020 inches from the casing endwall, and the mid-chord 

axial location is 0.400 inches upstream of the rotor blade stacking line. Corresponding contours of 

relative Mach number at the rotor C-grid downstream boundary, located nominally 0.400 inches 

downstream of the rotor trailing edge (0.300 at hub to 0.500 at tip), are shown in Figure 23, and 

contours of various flow properties at the rotor exit bypass H-grid downstream boundary, which is 

the downstream mixing plane located nominally 0.350 inches upstream of the FEGV leading edge 

(0.250 at hub to 0.450 at tip), are shown in Figures 24a and 24b. Figure 24a shows relative and 

absolute (stationary system) Mach number contours, and Figure 24b shows entropy and turbulence 

kinetic energy contours. Note that 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 operation at the SLTO operating 

point are shown in Figures 25a and 25b. Figure 25a shows a vane section at about 87 percent span 

from the hub, and Figure 25b shows vane sections at about 52 and 15 percent span (see red 

streamlines in Figure 7). Corresponding Mach number contours at the FEGV C-grid downstream 

boundary and exit H-grid downstream boundary, respectively, are shown in Figures 26 and 27.

Bypass strut blade-to-blade absolute Mach number contours for a location near midspan 

(see middle red streamline in Figure 7), for operation at the SLTO operating point, are shown in 

Figure 28. Corresponding Mach number contours at the strut grid downstream boundary are shown 

in Figure 29. Mach contours for the core-strut downstream boundary are included in Figure 29.

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

Figure 30, with an enlarged view of the fan region shown in Figure 31. 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 32a and 

32b, and various flow property contours at the rotor downstream mixing plane are shown in 

Figures 33a and 33b. Flow field contours for the FEGV at cutback operation are not shown, but are 

in general aerodynamically similar to those at SLTO.

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

in Figure 34, and Figure 35 shows an enlarged view of the fan region. Corresponding rotor bladeto-blade relative Mach number contours are shown in Figures 36a and 36b, and various flow 

property contours at the rotor downstream mixing plane are shown in Figures 37a and 37b. 

Absolute Mach number contours for the FEGV are shown in Figures 38a and 38b, and in 

Figures 39 and 40. The computed FEGV flow field, like that for cutback operation, is 

aerodynamically similar to the SLTO solution, at least generally. Notice in Figure 38a, however, 

the flow separation from the vane pressure side for the near tip section. This separation, caused by 

NASA/CR—2014-218131 9
a locally negative vane incidence, is not aerodynamically critical, but does increase the vane totalpressure loss slightly. A similar, but less severe separation is also present in the FEGV solution for 

the cutback operating point, while for the SLTO operating point (see Figure 25a) the separation is 

relatively minor. Nevertheless, despite this non-conformity and other differences, the solutions can 

be considered aerodynamically similar overall, showing that the FEGV flow field scales more-orless with flow rate for all simulated operating points on the fan operating line.

Finally, by comparing Figure 35 to Figure 19, it can be seen that the hub flow separation at 

the core duct entrance is larger for the approach operating point (also for the cutback operating 

point; compare Figure 31 to Figure 19) than for the SLTO operating point. This difference, which 

was alluded to earlier in the discussion of SLTO results, can also be seen in more detail, as well as 

from a more three-dimensional perspective, by comparing Figure 40 to Figure 27. These latter 

figures allow the non-averaged flow fields at the FEGV downstream mixing plane to be contrasted.

Concluding Remarks

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

only fan rotational speed and nozzle plug position 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. The comparison of 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 

flow separations or related secondary flow problems, with the exception of the hub boundary layer 

at the core duct entrance. At that location a significant flow separation is present. The region of 

local flow recirculation extends through a mixing plane, however, which for the particular mixingplane model used is now known to exaggerate the recirculation. In any case, the flow separation 

has relatively little impact on the computed rotor and FEGV flow fields.

NASA/CR—2014-218131 10
References

1. Kontos, Karen, Weir, Don, and Ross, David, ìQuiet High Speed Fan II (QHSF II) Final 

Report,î NASA CR-2012-217451, March 2012.

2. 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).

3. 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.

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

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

5. 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.

6. 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).

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

Documentation, Version 300, August 2003.

8. 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).

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

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

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

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

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

Third edition, 2006.

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

Documentation,î Version 300, July, 2003.

NASA/CR—2014-218131 11
Figure 1b: Enlarged Partial Schematic Drawing of the QHS Fan System

Figure 1a: Schematic Drawing of the QHS Fan System

-20 -10 0 10 20 30

Axial Coordinate, x (inches)

0

10

Radial Coordinate, 

r (inches)

Inlet Plug

Nacelle

Nozzle

-10 -5 0 5 10 15 20 25

Axial Coordinate, x (inches)

0

5

10

15

Radial Coordinate, 

r (inches)

FEGV

(50) Bypass Duct

Spinner

Fan

Rotor

(22)

Bypass

 Struts

 (10)

 Core

Struts

 (10)

Core Duct

NASA/CR—2014-218131 12
Figure 3: Chordwise Distributions of Rotor Blade Tip Clearance

Figure 2: Nacelle Trailing Edge Trim Configurations

34.5 35.0 35.5 36.0 36.5 37.0

Axial Coordinate, x (inches)

11.0

11.5

12.0

Radial Coordinate, 

r (inches)

trim-A

trim-1

trim-0

0 20 40 60 80 100

Meridional Chordwise Location (percent)

25

30

35

40

45

Rotor Tip Clearance Gap Height (mils)

14,060

9,840

10,935

12,500

13,280

Rotational

 Speed

 (rpm)

NASA/CR—2014-218131 13
Figure 4: Two-Dimensional Axisymmetric CFD Grid Blocks (Entire Computational Domain)

Figure 5a: Enlarged Inlet, Core and Bypass Duct, Nozzle, and Nacelle Grid Blocks

NASA/CR—2014-218131 14
Figure 5b: Enlarged Inlet Grid Block

Figure 5c: Enlarged Duct and Nozzle Grid Blocks

NASA/CR—2014-218131 15
Figure 6: Magnified Nacelle Leading and Trailing Edge Grids

Figure 7: Meridional Location and Extent of Three-Dimensional CFD Grid Blocks

Rotor

Inlet

H-Grid

Rotor

C-Grid

FEGV

C-Grid

Rotor Exit

H-Grid

Bypass

Strut

H-Grid

Core Strut

H-Grid FEGV Exit

Core

H-Grid

FEGV Exit

Bypass

H-Grid

NASA/CR—2014-218131 16
Figure 8: Meridional View of Three-Dimensional CFD Grid Blocks (Mid-Pitch Location)

NASA/CR—2014-218131 17
Figure 9: Three-Dimensional View of Rotor Grid

NASA/CR—2014-218131 18
Figure 10: Three-Dimensional View of FEGV Grid

NASA/CR—2014-218131 19
Figure 11: Rotor Grid at Three Spanwise Locations

Rotor Near Mid

Rotor Near Hub

Rotor Near Tip

NASA/CR—2014-218131 20
Figure 12: Rotor Grid at Block Downstream Boundaries

Rotor Exit

H-Grid

Downstream

Boundary

Rotor

C-Grid

Downstream

Boundary

Tip

Hub

NASA/CR—2014-218131 21
Figure 13: Rotor Tip Clearance Grid

Trailing Edge

Leading Edge

A ñ A

Blade Tip

Clearance

Gap

A

A

Rotor Tip

NASA/CR—2014-218131 22
Figure 14: FEGV Grid at Near Midspan Location

FEGV Near Midspan

Leading Edge Trailing Edge

NASA/CR—2014-218131 23
Figure 15: FEGV Grid at Block Downstream Boundaries

Tip

Hub

FEGV

C-Grid

Downstream

Boundary

FEGV Exit

H-Grid

Downstream

Boundary

Core

Bypass

NASA/CR—2014-218131 24
Figure 16: Core-Strut Grid at Near Midspan Location

Core Strut Near Midspan

Leading Edge Trailing Edge

NASA/CR—2014-218131 25
Figure 17: Bypass- and Core-Strut Grids at Block Downstream Boundaries

Tip

Hub

Bypass-Strut and Core-Strut H-Grid

Downstream Boundaries

Core

Bypass

NASA/CR—2014-218131 26
Figure 18: 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—2014-218131 27
Figure 19: 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—2014-218131 28
Figure 20: 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—2014-218131 29
Figure 21a: Computed Rotor Flow Field for the SLTO Operating Point;

Blade-to-Blade Relative Mach Number Contours

Rotor Near Tip

(87% Span)

Rotation

Flow

SLTO Operating Point

0.0 0.5 1.0 1.5

Relative Mach Number

NASA/CR—2014-218131 30
Figure 21b: Computed Rotor Flow Field for the SLTO Operating Point;

Blade-to-Blade Relative Mach Number Contours

Rotor Near Mid

(52% Span)

Rotor Near Hub

(15% Span)

NASA/CR—2014-218131 31
Figure 22: Computed Rotor Tip Flow Field for the SLTO Operating Point;

Relative Mach Number Contours (Legend in Figure 21a)

Mid-Chord

Axial Location

Mid-Clearance-Gap

Spanwise Location

Rotor Tip Clearance

Rotation

NASA/CR—2014-218131 32
Figure 23: 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—2014-218131 33
Figure 24a: 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

0.8

NASA/CR—2014-218131 34
Figure 24b: Computed Rotor Flow Field for the SLTO Operating Point; Downstream

Mixing Plane; Entropy and Turbulence Kinetic Energy Contours

Turbulence

Kinetic

Energy

0.000

0.008

0.002

0.004

0.006

Entropy

0.0

0.2

0.3

0.1

Rotor Exit

H-Grid

Downstream

Boundary

Rotation

0.4

0.5

NASA/CR—2014-218131 35
Figure 25a: Computed FEGV Flow Field for the SLTO Operating Point;

Blade-to-Blade Mach Number Contours

FEGV Near Tip

(87% Span)

Flow

SLTO Operating Point

0.0 0.4 0.8 1.0

Mach Number

0.2 0.6 1.2

NASA/CR—2014-218131 36
Figure 25b: Computed FEGV Flow Field for the SLTO Operating Point;

Blade-to-Blade Mach Number Contours

FEGV Near Mid

(52% Span)

FEGV Near Hub

(15% Span)

Flow

NASA/CR—2014-218131 37
Figure 26: Computed FEGV Flow Field for the SLTO Operating Point; Downstream

Boundary of C-Grid; Mach Number Contours

FEGV C-Grid

Downstream

Boundary

0.0

0.4

0.6

Mach

Number

0.2

NASA/CR—2014-218131 38
Figure 27: Computed FEGV Flow Field for the SLTO Operating Point;

Downstream Mixing Planes; Mach Number Contours

FEGV Exit

H-Grid

Downstream

Boundary

Bypass

Core

0.0

0.4

0.6

0.2

Mach

Number

NASA/CR—2014-218131 39
Figure 28: Computed Bypass Strut Flow Field for the SLTO Operating Point;

Blade-to-Blade Mach Number Contours

Bypass Strut

Near Midspan

Flow

SLTO Operating Point

0.0

0.4

0.6

Mach

Number

0.2

NASA/CR—2014-218131 40
Figure 29: Computed Bypass and Core Strut Flow Fields for the SLTO Operating Point;

Downstream Mixing Planes; Mach Number Contours

Bypass-Strut and Core-Strut H-Grid

Downstream Boundaries

Core

Bypass

0.0 0.2 0.3 0.5

Mach Number

0.1 0.4

NASA/CR—2014-218131 41
Figure 30: Computed Fan System Flow Field for the Cutback Operating Point;

Pitchwise-Averaged Flow; Mach Number Contours

0.0 0.2 0.4 0.6 0.8

Mach Number

Cutback Operating Point

NASA/CR—2014-218131 42
Figure 31: 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—2014-218131 43
Figure 32a: Computed Rotor Flow Field for the Cutback Operating Point;

Blade-to-Blade Relative Mach Number Contours

Rotor Near Tip

(87% Span)

Rotation

Flow

Cutback Operating Point

0.0 0.5 1.0 1.5

Relative Mach Number

NASA/CR—2014-218131 44
Figure 32b: Computed Rotor Flow Field for the Cutback Operating Point;

Blade-to-Blade Relative Mach Number Contours

Rotor Near Mid

(52% Span)

Rotor Near Hub

(15% Span)

NASA/CR—2014-218131 45
Figure 33a: 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

1.0

0.8

NASA/CR—2014-218131 46
Figure 33b: Computed Rotor Flow Field for the Cutback Operating Point; Downstream

Mixing Plane; Entropy and Turbulence Kinetic Energy Contours

Turbulence

Kinetic

Energy

0.000

0.005

0.002

0.003

0.004

Entropy

0.0

0.2

0.3

0.1 Rotor Exit

H-Grid

Downstream

Boundary

Rotation

0.4

0.001

NASA/CR—2014-218131 47
Figure 34: Computed Fan System Flow Field for the Approach Operating Point;

Pitchwise-Averaged Flow; Mach Number Contours

0.0 0.2 0.4 0.6

Mach Number

Approach Operating Point

0.8

NASA/CR—2014-218131 48
Figure 35: Computed Fan Blade Row Flow Fields for the Approach Operating Point;

Pitchwise-Averaged Flow; Mach Number Contours

0.0 0.2 0.4 0.6

Mach Number

0.8

NASA/CR—2014-218131 49
Figure 36a: Computed Rotor Flow Field for the Approach Operating Point;

Blade-to-Blade Relative Mach Number Contours

Rotor Near Tip

(87% Span)

Rotation

Flow

Approach Operating Point

0.0 0.5 1.0 1.5

Relative Mach Number

NASA/CR—2014-218131 50
Figure 36b: Computed Rotor Flow Field for the Approach Operating Point;

Blade-to-Blade Relative Mach Number Contours

Rotor Near Mid

(52% Span)

Rotor Near Hub

(15% Span)

NASA/CR—2014-218131 51
Figure 37a: 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

0.4

0.6

Rotor Exit

H-Grid

Downstream

Boundary

Relative

Mach

Number

Rotation

0.0

0.8

0.2

0.4

0.6

NASA/CR—2014-218131 52
Figure 37b: Computed Rotor Flow Field for the Approach Operating Point; Downstream

Mixing Plane; Entropy and Turbulence Kinetic Energy Contours

Turbulence

Kinetic

Energy

0.000

0.004

0.001

0.002

0.003

Entropy

0.00

0.10

0.15

0.05

Rotor Exit

H-Grid

Downstream

Boundary

Rotation

0.20

0.25

NASA/CR—2014-218131 53
Figure 38a: Computed FEGV Flow Field for the Approach Operating Point;

Blade-to-Blade Mach Number Contours

FEGV Near Tip

(87% Span)

Flow

Approach Operating Point

0.0 0.4 0.8

Mach Number

0.2 0.6

NASA/CR—2014-218131 54
Figure 38b: Computed FEGV Flow Field for the Approach Operating Point;

Blade-to-Blade Mach Number Contours

FEGV Near Mid

(52% Span)

FEGV Near Hub

(15% Span)

Flow

NASA/CR—2014-218131 55
Figure 39: Computed FEGV Flow Field for the Approach Operating Point; Downstream

Boundary of C-Grid; Mach Number Contours

FEGV C-Grid

Downstream

Boundary

0.0

0.4

0.5

Mach

Number

0.2

0.3

0.1

NASA/CR—2014-218131 56
Figure 40: Computed FEGV Flow Field for the Approach Operating Point;

Downstream Mixing Planes; Mach Number Contours

FEGV Exit

H-Grid

Downstream

Boundary

Bypass

Core

0.0

0.4

0.5

Mach

Number

0.2

0.3

0.1

NASA/CR—2014-218131 57
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4. TITLE AND SUBTITLE

Computational Aerodynamic Simulations of a 1484 ft/sec Tip Speed Quiet High-Speed

Fan System Model for Acoustic Methods Assessment and Development

5a. CONTRACT NUMBER

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13. SUPPLEMENTARY NOTES

14. ABSTRACT

Computational Aerodynamic simulations of a 1484 ft/sec tip speed quiet high-speed 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 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 includes a core duct and a bypass duct that merge upstream of the fan system nozzle. As a result, only

fan rotational speed and the system bypass ratio, set by means of a translating nozzle plug, 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. 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 computed flow

fields reveals no excessive or critical boundary layer separations or related secondary-flow problems, with the exception of the hub boundary layer at the core duct entrance. At

that location a significant flow separation is present. The region of local flow recirculation extends through a mixing plane, however, which for the particular mixing-plane

model used is now known to exaggerate the recirculation. In any case, the flow separation has relatively little impact on the computed rotor and FEGV flow fields.

15. SUBJECT TERMS

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

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