Mark L. Celestina and Kenneth L. Suder

Glenn Research Center, Cleveland, Ohio

Sameer Kulkarni

Case Western Reserve University, Cleveland, Ohio

Results of an Advanced Fan Stage Operating 

Over a Wide Range of Speed and Bypass Ratio

Part 2: Comparison of CFD and Experimental Results

NASA/TM—2010-216769/PART2

October 2010

GT2010–23386
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Mark L. Celestina and Kenneth L. Suder

Glenn Research Center, Cleveland, Ohio

Sameer Kulkarni

Case Western Reserve University, Cleveland, Ohio

Results of an Advanced Fan Stage Operating 

Over a Wide Range of Speed and Bypass Ratio

Part 2: Comparison of CFD and Experimental Results

NASA/TM—2010-216769/PART2

October 2010

GT2010–23386

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

Turbo Expo 2010

sponsored by the American Society of Mechanical Engineers (ASME)

Glasgow, Scotland, United Kingdom, June 14–18, 2010
Acknowledgments

The authors would like to acknowledge the helpful discussions with Dr. John Adamczyk (NASA Distinguished Research Associate) 

for his insights and understanding of turbomachinery as well as the contributions of Mr. Timothy Beach whose turbomachinery 

gridding experience with his mesh code was invaluable to this work.

Available from

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This work was sponsored by the Fundamental Aeronautics Program 

at the NASA Glenn Research Center.

Level of Review: This material has been technically reviewed by technical management. 
NASA/TM—2010-216769/PART2 1

Results of an Advanced Fan Stage Operating 

Over a Wide Range of Speed and Bypass Ratio

Part 2: Comparison of CFD and Experimental Results

Mark L. Celestina and Kenneth L. Suder

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Sameer Kulkarni

Case Western Reserve University

Cleveland, Ohio 44106

Abstract

NASA and GE teamed to design and build a 57 percent

engine scaled fan stage for a Mach 4 variable cycle 

turbofan/ramjet engine for access to space with multipoint 

operations. This fan stage was tested in NASA’s transonic 

compressor facility. The objectives of this test were to assess 

the aerodynamic and aero mechanic performance and 

operability characteristics of the fan stage over the entire range 

of engine operation including: 1) sea level static take-off; 

2) transition over large swings in fan bypass ratio; 3) transition 

from turbofan to ramjet; and 4) fan wind-milling operation at 

high Mach flight conditions. 

This paper will focus on an assessment of APNASA, a 

multistage turbomachinery analysis code developed by NASA, 

to predict the fan stage performance and operability over a 

wide range of speeds (37 to 100 percent) and bypass ratios.

Introduction

The research objectives were to assess the capability of

state-of-the-art (SOA) design and analysis tools used for 

conventional subsonic turbomachinery to design and predict 

the performance and operability of an advanced fan stage 

designed to meet the requirements for the first stage of a twostage-to-orbit hypersonic vehicle. The ultimate goal was to 

have confidence in the tools to design and analyze these 

advanced Turbine Based Combined Cycle (TBCC) 

configurations to meet future mission requirements. The 

metrics for the CFD analysis was to determine fan stage 

performance in terms of adiabatic efficiency (at a design 

pressure ratio and mass flow) and fan stage operability in 

terms of stall margin (over the operating line).

The approach taken to assess the SOA design and analysis 

tools was to: 1) design and predict the performance and 

operability of a relevant TBCC engine fan stage; 

2) experimentally map the aerodynamic and aeromechanic 

performance and stability limits over the operating range; 

3) compare measured results to pretest CFD predictions; and 

4) assess the strengths and weaknesses of the SOA tools.

In this paper, we will describe the lessons learned in a 

pretest CFD analysis and results of comparing the post-test 

CFD analysis with experimental data then wrap-up with an 

overall comparison of the design with the CFD and 

experimental results. The CFD tool used in this paper is 

APNASA (Ref. 1).

Nomenclature

MF Mass Flow

NS Near Stall Point

OP Operating Point

PR Pressure Ratio

SM Stall Margin

TR Temperature Ratio

RTA Fan Rig

The GE57 fan rig was designed as a 57 percent linear scale 

of the proposed demonstrator engine. It was designed to 

demonstrate the feasibility of the high-pressure ratio fan 

across the required broad operating range. The fan diameter is 

55.88 cm (22 in.) and at 100 percent speed develops a design 

pressure ratio of 2.46 at an inlet corrected flow rate of 

37.19 kg/s (82 lbm/s). The stage adiabatic efficiency at this 

point was projected to be 85.7 percent and stall margin was 

projected to be 11 percent. The fan rotational speed at 

100 percent speed is 17,280 rpm with a corresponding tip 

speed of about 506 m/s (1660 ft/s). A variable fan OGV was 

designed for this rig for high Mach flight conditions.

Following the OGV is the split flowpath with struts that 

extend through both flowpaths. The lower path directs flow 

through the core of the engine and the upper path bypasses the 

core. Two liners over the rotor were used in the experiment, a 

grooved casing and a smooth casing. The CFD was only run 

for the smooth liner. For more details about the design, see

(Ref. 2). For information on the instrumentation survey planes, 

see (Ref. 3).
NASA/TM—2010-216769/PART2 2

Figure 1.—Fan stage components and flowpath.

Figure 2.—Axisymmetric view of simulation grid.

APNASA Code

APNASA is a multistage analysis code which provides the 

ability to analyze the complete fan stage consisting of the fan 

rotor, outlet guide vane (OGV), and splitter/strut mid-frame 

assembly (see Fig. 1). This approach is unique from ‘typical’ 

single passage codes that solve for one blade row at a time by 

systematically marching through the compression system 

blade row by blade row utilizing averaged or mixed-out 

solutions from the upstream blade row as input to the 

downstream blade row. APNASA solves for all blade rows 

under consideration which, in this case, includes the fan rotor, 

OGV and splitter/strut mid-frame assembly. The APNASA 

fan rotor solution includes body force terms to represent the 

effect of the downstream OGV and splitter/strut. Similarly the 

OGV solution includes body force terms to account for the 

upstream fan rotor and the downstream strut/splitter, 

respectively. The numerical procedure solves the threedimensional Reynolds averaged Navier-Stokes and energy 

equations using the four step Runge-Kutta scheme and is 

described by Adamczyk, et al. (Ref. 3). The standard twoequation k–ε model was used to calculate the turbulent eddy 

viscosity coefficient as described by Shabbir et al. (Refs. 4 and 

5). Spalding’s formula (Ref. 6) was used to set the wall 

boundary conditions for the mean flow equations as well as 

the turbulence model. The inlet boundary conditions consisted 

of radial profiles of total pressure, total temperature, and radial 

and tangential flow angle which were deduced from the 

experimental data. The flow in the clearance gap was 

simulated using a model suggested by Kirtley et al. (Ref. 7)

and treats the clearance flow as an orifice flow with no loss in 

mass, momentum, or energy. The effect of the vena contracta, 

which occurs in orifice flow is accounted for by using a 

discharge coefficient, which makes the effective tip clearance 

gap smaller than the actual clearance. Discharge coefficients 

of 0.5 to 1.0 are typically used (Ref. 8), however, a discharge 

coefficient of 1.0 is used for all results presented as advocated 

by Van Zante, et al., (Ref. 9). Typically 1 to 4 cells are used to 

define the clearance region. For this paper, the computational 

grids allow the number of cells in the clearance gap to vary 

with operating speed. In these simulations the tip clearance 

gap increased as the rotor speed decreased consistent with the 

thermal and mechanical stresses on the rotor and casing. The 

assessment of the change in clearances with rotor speed and 

inlet temperatures was provided by the experimental data. The 

APNASA solver expects a unique mesh for the rotor fan, 

OGV and strut, each of which shares a common axisymmetric 

grid as shown in Figure 2. The computational grids used for 

the present calculations were generated using TGS, a 

turbomachinery grid code developed by Beach (Ref. 10). The 

grids used by APNASA are sheared H-type in the blade-toblade direction and incorporate a hyperbolic tangent stretching 

parameter to resolve the boundary layer and adequately define 

the flow regions. Similar stretching functions are chosen for 

the span-wise and chord- wise directions to enable adequate 

resolution of the blade leading and trailing edges and

clearance gaps. The blade-to-blade stretching is relaxed to 

uniform spacing upstream and downstream of the blade rows 

since the flow is assumed to be periodic in those regions.

The grids used in the simulations had 80 cells in the radial 

direction, 50 cells in the circumferential direction with a total 

of 373 cells placed between the inlet and exit boundaries of 

which 60 were along the chord lines of each blade row (fan 

rotor, OGV, mid-frame strut).

Pretest Lessons Learned

Prior to the experiment, APNASA was run at 80 to 

100 percent of rotor design speed. These pretest simulation 

results were compared to the experimental results of the fan 

stage characteristics at these speeds for the overall stage 

pressure ratio and adiabatic efficiency (see (Ref. 3)). In 

general, the simulations compared favorably with the data 

along the operating line both in terms of mass flow, pressure 

ratio, and to a lesser degree efficiency. However, a major 

concern at the time for the simulations was the inability of the 

code to predict the stall margin at 100 percent design speed.

At 80 percent speed, the stall margin was well predicted.

Analysis of the simulations as well as later comparisons with 

the experimental data led to the following issues:

INLET 

CORE FLOW

BYPASS FLOW

ROTOR

OGV

STRUT
NASA/TM—2010-216769/PART2 3

1. The pretest simulation was run at the same bypass ratio as 

the data only along the operating line but at all other 

operating conditions along a speedline, the bypass ratio for 

the data and simulations were significantly different.

2. The measured rotor tip clearance was substantially lower 

than the clearance assumed by the simulations across the 

entire speed range. For example, the measured clearance at 

mid-chord of the rotor at 100 percent design speed was 

0.30 mm (0.0012 in.) compared to the design intent of 

0.51 mm (0.0020 in.) at design speed used in the 

computation.

3. The blade geometry was not properly hot-shape corrected 

resulting in the wrong blade geometry being simulated 

which can have a large effect on fan performance.

For the post-test assessment, the following adjustments 

were made to the grid and computational approach to 

circumvent these problems:

1. A mass flow boundary condition (Ref. 11) was used in the 

core flow which allowed the simulation to run with a 

boundary condition (in terms of bypass ratio) close to that 

at which the experimental data was acquired. This change

allowed the bypass ratio of the simulation to better match 

the bypass ratio set in the experiment.

2. The simulation used the measured average tip clearance 

corresponding to each fan rotor speed. The average tip 

clearance was obtained by averaging the dynamic 

measurements of tip clearance for each blade made at the 

mid-chord region of the rotor blade at four different 

circumferential locations around the casing.

3. The blade geometry was corrected for rotational, 

temperature, and pressure loading corresponding to: 1) the 

blade cold shape, 2) hot shape corrected at 80 percent

speed, and 3) hot shape geometry at 100 percent speed.

Post-Test simulations were performed with only the 

changes described above. This ensured that the conditions 

used in the simulations were similar to those of the 

experiment. The intent of these changes was to provide for a 

realistic assessment of the codes ability to predict the fan stage 

performance and operability.

Posttest APNASA Results and 

Comparison to Data

Post-Test simulations were performed at 100, 90, 80, 60, 

and 37 percent of fan rotor design speed. The hot shape 

corrected for 100 percent speed geometry was used for the 

90 percent speed line simulations. The cold shape geometry 

was used for the 60 and 37 percent speed line simulations. The 

tip clearance over the rotor was assumed constant over the 

chord. The tip clearance values used in the computations are 

presented in Table 1 with the ratio of clearance gap to tip 

chord. Comparisons between the data and simulations were 

made in terms of pressure ratio and efficiency for the entire 

operating range at each of these speedlines. The pressure and 

temperature ratio were area-averaged and the adiabatic 

efficiency calculated based on the value of those two 

parameters. Tabulated results at choke, the operating point and 

near stall are presented with stall margin at the above fan 

speeds. The stall margin, SM, is defined as

[ ]

[ ]OP

NS

PR MF

PR MF

SM =

where PR is the pressure ratio, MF is the massflow, NS is the 

near stall point and OP is the operating point. The tip casing 

static pressure is also compared to data and is presented at the 

operating point for each of the fan speeds. Axisymmetric 

profiles are presented at the operating point and at the near 

stall point and compared to data at 100, 80, and 37 percent of 

the design speed. After these comparisons are made, an overall

comparison is made of the data and the simulations, to the

design intent only along the operating line at these same 

speeds.

TABLE 1.—SIMULATION ROTOR TIP

CLEARANCE AT PERCENT SPEED

Speed

(%)

mm in. Gap/chord

(%)

100 0.30 0.012 0.26

90 .36 .014 .31

80 .41 .016 .35

60 .51 .020 .44

37 .66 .026 .57

Comparisons at 100 Percent Speed

Comparisons between the data and simulations at 100 percent

design speed for the stage pressure ratio and adiabatic 

efficiency are shown in Figure 3. The experimental data is 

shown in red and the analysis in blue. The blue line represents 

the averaged results using the simulation results across the 

entire span and the blue symbols represent the averaged results 

using only the radial locations consistent with the radial 

locations that the measurements were acquired. This color 

scheme and symbol representation persists throughout this 

paper. It is evident from the results depicted in Figure 3 that 

the CFD does a credible job of predicting the pressure ratio 

and efficiency over the entire range of back-pressures and not 

just at the operating point for 100 percent speed. A more 

quantitative comparison is provided in Table 2. The 

parameters of corrected mass flow, pressure ratio (PR), 

temperature ratio (TR), adiabatic efficiency, and stall margin 

are provided for the data and simulation for points along the 

speedline. These points are choke, the operating line, and near 

stall. The only valid comparison for the choke flow condition 

is to compare the mass flow for the data to that of the 

simulation because the fan stage characteristic is essentially 

vertical. No attempt was made in the simulation to match the 
NASA/TM—2010-216769/PART2 4

data at the lowest pressure ratio. However, it was felt 

important to report the maximum flow condition and compare 

to the data, especially at high rotational speeds. If there is 

good agreement in the choked flow between the data and the 

simulation, then there is a high degree of confidence that the 

geometry is being accurately modeled in the simulations. It is 

also evident that the agreement between the simulation and the 

data is much better at the operating line than at the choked 

flow and near stall conditions where the likelihood of 

separated and unsteady flow effects are much higher.

TABLE 2.—QUANTITATIVE COMPARISON OF SIMULATION 

RESULTS TO DATA AT 100 PERCENT SPEED

Speed (%) 100

Flow Choke Op NS

Data Mass flow

(kg/s)

37.459 37.318 34.827

APNASA 37.432 37.328 34.645

Data Pressure 

ratio (PR)

2.518 2.571

APNASA 2.486 2.557

Data Temperature 

ratio (TR)

1.354 1.383

APNASA 1.342 1.37

Data Efficiency 0.844 0.799

APNASA 0.865 0.828

Data Stall margin 

(%)

9.41

APNASA 10.82

Figure 3 indicates that the simulation is producing higher 

efficiencies than the data. Further comparisons between the 

data and simulation indicates that the fan rotor static pressure 

rise at the tip is being over-predicted by the simulation (see 

Fig. 4) which is indicative of not accurately capturing the fan 

endwall flow characteristics which could be due to not 

modeling the non-uniform distribution of the fan tip clearance 

from leading edge to trailing edge. It should be noted that in 

the experiment, the rotor tip rubbed in near the trailing edge at 

100 percent speed.

Profiles of total pressure and efficiency behind the OGV at

the operating point and near stall are shown in Figure 5. At the 

operating line, the simulation total pressure in Figure 5(a) 

exhibits a lower level of total pressure than the data but with a 

consistent shape. There also appears to be a higher level of 

total pressure near the hub in the simulation that is not seen in 

the data. The simulation efficiency overall in Figure 5(b) is 

higher than the data which was pointed out earlier but the 

shape follows the general trend of the data. Since the 

efficiency matches so well near the tip and the total pressure 

ratio is lower, the simulation overpredicts the measured total 

temperature.

At near stall in Figure 5(c), the total pressure level below 

midspan from the simulation is lower than experiment and is 

overpredicted in the upper 50 percent span. Similarly in 

Figure 5(d), the efficiency is underpredicted near the hub and 

overpredicted from 40 percent span outward. The wiggles in 

the profile below midspan indicate weak (separated) flow near 

the hub that is not evident in the experimental data.

Figure 3.—Stage pressure ratio and adiabatic efficiency versus

corrected inlet flow at 100 percent speed.

Figure 4.—Comparison of casing static pressure with data at 

100 percent speed at the operating point.

1.9

2

2.1

2.2

2.3

2.4

2.5

2.6

2.7

34.5 35.0 35.5 36.0 36.5 37.0 37.5

Total Pressure Ratio

Corrected Inlet Flow , kg/s

DATA

APNASA

0.6

0.65

0.7

0.75

0.8

0.85

0.9

34.5 35.0 35.5 36.0 36.5 37.0 37.5

Adiabatic Efficiency

Corrected Inlet Flow , kg/s

DATA

APNASA

0.5

0.7

0.9

1.1

1.3

1.5

1.7

1.9

2.1

-100 -50 0 50 100 150

Tip Static Pressure Ratio

Percent Rotor Chord

APNASA

DATA
NASA/TM—2010-216769/PART2 5

 

 

Figure 5.—Profiles of total pressure ratio and adiabatic efficiency aft of the OGV at the operating point ((a) and

(b)) and near stall ((c) and (d)) at 100 percent speed.

Figure 6.—Stage pressure ratio and adiabatic efficiency versus corrected inlet flow at 90 percent speed.

Comparisons at 90 Percent Speed

Comparisons between the data and simulations at 

90 percent speed for the stage pressure ratio and adiabatic 

efficiency are shown in Figure 6. The simulation results at 

90 percent speed, similar to those at 100 percent speed, show 

good agreement across the operating range. Pressure ratio is 

well predicted at the operating line but the efficiency is still 

over-predicted by about 3 points at the operating point similar 

to the 100 percent speed results. Tabulated results for choke, 

the operating point and near stall are found in Table 3. Stall 

margin from the simulation was slightly lower than 

determined from the data.

0

20

40

60

80

100

2 2.1 2.2 2.3 2.4 2.5 2.6 2.7 2.8 2.9 3

Percent Span

a) Total Pressure Ratio

APNASA

DATA

0

20

40

60

80

100

0.6 0.65 0.7 0.75 0.8 0.85 0.9 0.95 1

Percent Span

b) Adiabatic Efficiency

APNASA

DATA

0

20

40

60

80

100

2 2.1 2.2 2.3 2.4 2.5 2.6 2.7 2.8 2.9 3

Percent Span

c) Total Pressure Ratio

APNASA

DATA

0

20

40

60

80

100

0.6 0.65 0.7 0.75 0.8 0.85 0.9 0.95 1

Percent Span

d) Adiabatic Efficiency

APNASA

DATA

1.7

1.8

1.9

2

2.1

2.2

2.3

28 29 30 31 32 33 34 35 36

Total Pressure Ratio

Corrected Inlet Flow, kg/s

DATA

APNASA

0.68

0.7

0.72

0.74

0.76

0.78

0.8

0.82

0.84

0.86

0.88

0.9

0.92

28 29 30 31 32 33 34 35 36

Adiabatic Efficiency

Corrected Inlet Flow, kg/s

DATA

APNASA
NASA/TM—2010-216769/PART2 6

TABLE 3.—QUANTITATIVE COMPARISON OF SIMULATION 

RESULTS VERSUS DATA AT 90 PERCENT SPEED

Speed (%) 90

Flow Choke Op NS

Data Mass flow

(kg/s)

35.084 33.178 29.410

APNASA 34.956 33.035 29.009

Data Pressure 

ratio (PR)

2.195 2.202

APNASA 2.24 2.161

Data Temperature 

ratio (TR)

1.292 1.316

APNASA 1.29 1.304

Data Efficiency 0.858 0.793

APNASA 0.89 0.808

Data Stall margin 

(%)

13.17

APNASA 9.86

Comparisons at 80 Percent Speed

Comparisons between the data and simulations at 

80 percent design speed for the stage pressure ratio and 

adiabatic efficiency are shown in Figure 7 with the 

quantitative comparisons tabulated in Table 4. At this speed, 

the agreement in total pressure is very good and differs from 

the data by less than 1 percent even at near stall. The 

efficiency calculation is overpredicted by less than 3 points at 

the operating point and slightly more than 3 points at near 

stall. The stall margin from the simulation agrees very well 

with the data. Figure 8 shows a comparison of the casing static 

pressure between the simulation and the data at the operating 

line. It shows that the pressure at the casing from the 

simulation agrees very well with the data over the rotor chord 

indicating that the shock location between the data and the 

simulation are practically identical. Profiles of total pressure 

ratio and adiabatic efficiency on the operating line, (a) and (b),

and near stall, (c) and (d) are shown in Figure 9. Similar to the 

100 percent speed profiles, the simulation profiles for total 

pressure ratio and adiabatic efficiency agree quite well with 

data at the operating line. Near the hub, however, the 

simulation profile indicates a stronger hub flow than the 

experiment. This stronger hub flow persists to near stall as 

indicated by the level of higher total pressure near the hub.

TABLE 4.—QUANTITATIVE COMPARISON OF SIMULATION 

RESULTS VERSUS DATA AT 80 PERCENT SPEED

Speed (%) 80

Flow Choke Op NS

Data Mass flow

(kg/s)

30.529 29.233 24.810

APNASA 30.281 28.206 25.014

Data Pressure 

ratio (PR)

1.855 1.869

APNASA 1.858 1.886

Data Temperature 

ratio (TR)

1.224 1.244

APNASA 1.222 1.239

Data Efficiency 0.857 0.797

APNASA 0.872 0.831

Data Stall margin 

(%)

14.66

APNASA 14.46

Figure 7.—Stage pressure ratio and adiabatic efficiency 

versus corrected inlet flow at 80 percent speed.

Figure 8.—Comparison of casing static pressure with 

data at 80 percent speed at operating point.

1.4

1.5

1.6

1.7

1.8

1.9

2

24 25 26 27 28 29 30 31

Total Pressure Ratio

Corrected Inlet Flow, kg/s

DATA

APNASA

0.66

0.68

0.7

0.72

0.74

0.76

0.78

0.8

0.82

0.84

0.86

0.88

0.9

24 25 26 27 28 29 30 31

Adiabatic Efficiency

Corrected Inlet Flow, kg/s

DATA

APNASA

0.6

0.8

1

1.2

1.4

1.6

-100 -50 0 50 100 150

Tip Static Pressure Ratio

Percent Rotor Chord

APNASA

DATA
NASA/TM—2010-216769/PART2 7

 

 Figure 9.—Profiles of total pressure ratio and adiabatic efficiency aft of the OGV at the operating point ((a) and (b)) and 

near stall ((c) and (d)) at 80 percent speed.

Figure 10.—Stage pressure ratio and adiabatic efficiency versus corrected inlet flow at 60 percent speed.

Comparisons at 60 Percent Speed

Comparisons between the data and simulations at 60 percent 

design speed for the stage pressure ratio and adiabatic efficiency 

are shown in Figure 10 with the results tabulated in Table 5. At 

this speed, the agreement in total pressure is very good and 

differs from the data by less than 1 percent even at near stall. 

The efficiency calculation is again overpredicted by less than 

3 points but is now nearly within the uncertainty measurements 

of the data. The stall margin from the simulation is 3.31 points 

lower than the data.

0

20

40

60

80

100

1.7 1.75 1.8 1.85 1.9 1.95 2

Percent Span

a) Total Pressure Ratio

APNASA

DATA

0

20

40

60

80

100

0.6 0.65 0.7 0.75 0.8 0.85 0.9 0.95 1

Percent Span

b) Adiabatic Efficiency

APNASA

DATA

0

20

40

60

80

100

1.6 1.7 1.8 1.9 2

Percent Span

c) Total Pressure Ratio

APNASA

DATA

0

20

40

60

80

100

0.6 0.7 0.8 0.9 1

Percent Span

d) Adiabatic Efficiency

APNASA

DATA

1.05

1.1

1.15

1.2

1.25

1.3

1.35

1.4

1.45

18 19 20 21 22 23

Total Pressure Ratio

Corrected Inlet Flow, kg/s

DATA

APNASA

0.64

0.66

0.68

0.70

0.72

0.74

0.76

0.78

0.80

0.82

0.84

0.86

0.88

18 19 20 21 22 23

Adiabatic Efficiency

Corrected Inlet Flow, kg/s

DATA

APNASA
NASA/TM—2010-216769/PART2 8

TABLE 5.—QUANTITATIVE COMPARISON OF SIMULATION 

RESULTS VERSUS DATA AT 60 PERCENT SPEED.

Speed (%) 60

Flow Choke Op NS

Data Mass flow

(kg/s)

22.800 22.091 18.255

APNASA 22.623 22.051 18.518

Data Pressure 

ratio (PR)

1.329 1.417

APNASA 1.324 1.404

Data Temperature 

ratio (TR)

1.107 1.128

APNASA 1.108 1.122

Data Efficiency 0.793 0.817

APNASA 0.77 0.836

Data Stall margin 

(%)

29.03

APNASA 26.27

Comparisons at 37 Percent Speed

Comparisons between the data and simulations at 37 percent 

design speed for the stage pressure ratio and adiabatic efficiency 

are shown in Figure 11 with the results tabulated in Table 6. At 

this speed, the agreement in total pressure is very good over 

most of the speedline, but is underpredicted at flow less than 

12 kg/s and more than 16 kg/s. The efficiency, however, tracks 

the data very well with a slight overprediction near stall. The 

stall margin from the simulation is a little more than 2.5 points 

lower than the data. Figure 12 shows a comparison of the casing 

static pressure between the simulation and the data at the 

operating line. The pressure at the casing from the simulation 

agrees very well with the data over the over front half of the 

rotor chord and higher than the data aft of mid-chord. This overprediction may be due to the tighter measured clearance in the 

back end of the rotor which was reported but not simulated. 

Profiles of stage total pressure ratio and adiabatic efficiency on 

the operating line and near stall are shown in Figure 13. The 

simulation profiles for total pressure ratio agree quite well with 

data at the operating line above 40 percent span. Near the hub, 

the simulation shows a rather large total pressure deficit up to 

about 20 percent which is not seen in the data. At near stall, the 

simulation exhibits a similar profile to the data but at a slightly 

lower level. The efficiency calculation from the data has a large 

degree of uncertainty because the total temperature rise at 

37 percent design speed is quite low. See (Ref. 3).

TABLE 6.—QUANTITATIVE COMPARISON OF SIMULATION 

RESULTS VERSUS DATA AT 37 PERCENT SPEED

Speed (%) 37

Flow Choke Op NS

Data Mass flow

(kg/s)

18.767 17.349 13.311

APNASA 18.236 17.355 10.552

Data Pressure 

ratio (PR)

1.03 1.139

APNASA 1.025 1.135

Data Temperature 

ratio (TR)

1.023 1.048

APNASA 1.025 1.044

Data Efficiency 0.372 0.791

APNASA 0.412 0.838

Data Stall margin 

(%)

86.06

APNASA 82.11

Figure 11.—Stage pressure ratio and adiabatic efficiency 

versus corrected inlet flow at 37 percent speed.

Figure 12.—Comparison of casing static pressure with 

data at operating point at 37 percent speed.

0.95

1

1.05

1.1

1.15

1.2

10 11 12 13 14 15 16 17 18 19

Total Pressure Ratio

Corrected Inlet Flow, kg/s

DATA

APNASA

0.4

0.5

0.6

0.7

0.8

0.9

1

10 11 12 13 14 15 16 17 18 19

Adiabatic Efficiency

Corrected Inlet Flow, kg/s

DATA

APNASA

0.86

0.88

0.9

0.92

0.94

0.96

0.98

1

1.02

-100 -50 0 50 100 150

Tip Static Pressure Ratio

Percent Rotor Chord

APNASA

DATA
NASA/TM—2010-216769/PART2 9

 

 

Figure 13.—Profiles of total pressure ratio and adiabatic efficiency aft of the OGV for the operating point ((a) and (b)) 

and near stall ((c) and (d)) at 37 percent speed.

Overall Comparisons Along the Design 

Operating Line

Some overall parameters were compared to design intent 

along the operating line. Figure 14 plots the total pressure ratio 

along the entire operating line from 37 to 100 percent fan 

speed for the results obtained from the design intent, i.e., GE’s 

design cycle deck, and compared to the experimental data and 

APNASA. 

The percent rotational speed of the fan is plotted along the 

abscissa, with the stage total pressure ratio plotted along the 

ordinate. APNASA tracks the data (red circles) very well, 

however, note how well the design intent also tracks the data. 

Figure 15 compares the corrected inlet mass flow with percent 

speed along the operating line. Again, it is difficult to 

distinguish between APNASA and the experimental data on 

this scale. The symbol for the data has been exaggerated for 

clarity. Figure 16 compares adiabatic efficiency with percent 

design speed also along the operating line. As was mentioned 

earlier in the comparisons, APNASA overpredicts the 

efficiency. As the percent speed drops to below 80 percent, the

Figure 14.—Comparison of stage pressure ratio 

versus percent speed along the operating line.

0

20

40

60

80

100

0.85 0.9 0.95 1 1.05 1.1 1.15

Percent Span

a) Total Pressure Ratio

APNASA

DATA

0

20

40

60

80

100

0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1

Percent Span

b) Adiabatic Efficiency

APNASA

DATA

0

20

40

60

80

100

1.10 1.12 1.14 1.16 1.18 1.20

Percent Span

c) Total Pressure Ratio

APNASA

DATA

0

20

40

60

80

100

0.6 0.65 0.7 0.75 0.8 0.85 0.9 0.95 1

Percent Span

d) Adiabatic Efficiency

APNASA

DATA

1

1.2

1.4

1.6

1.8

2

2.2

2.4

2.6

2.8

3

30% 40% 50% 60% 70% 80% 90% 100% 110%

Stage Pressure Ratio

Percent of Rotor Design Speed

DESIGN

DATA

APNASA
NASA/TM—2010-216769/PART2 10

difference in efficiency decreases. This overprediction of 

efficiency has been shown reported in the literature but the 

cause for this is not known. Figure 17 compares stall margin 

versus percent design speed. Again, the symbol for the 

experimental data has been exaggerated for clarity. Overall, 

APNASA was able to track the stall margin computed from 

the experimental data quite well which points to the tools 

suitability to analyze similar concepts.

Figure 15.—Comparison of corrected inlet mass flow 

versus percent speed along the operating line.

Figure 16.—Comparison of adiabatic efficiency versus

percent design speed along the operating line.

Figure 17.—Comparison of stall margin versus percent 

design speed.

Summary and Concluding Remarks

The RTA fan rig was simulated by APNASA for a wide 

operating range of speeds (37 to 100 percent) following some 

lessons learned from a pretest prediction of the same geometry.

These lessons included the importance of having the correct hot

shape, setting the proper tip clearance for the rotor simulations 

and understanding what boundary conditions are necessary to 

obtain good simulation results. These simulations were 

compared to experimental data taken at the NASA Glenn 

Research Center in their High Speed Compressor facility. The 

comparisons showed that the stage pressure ratio compared 

quite well to the data with the efficiency overpredicted by less 

than 3 points. Profiles of total pressure ratio and adiabatic 

efficiency were compared to data at both the operating line and 

near stall. The simulation profiles, for the most part, compared 

well to the data. In the near stall case at design speed, APNASA 

predicted a weaker hub flow than the data which will be looked 

at more carefully. APNASA was able to compute the flow at all 

the range of speeds required as well as obtain a fan stall margin 

that compared quite well with the data. This work shows that 

APNASA can be used to help design and analyze similar future 

fan stage concepts.

References

1. Adamczyk, J.J., Celestina, M.L., Beach, T.A. and Barnet, M., 

“Simulation of Three-dimensional Viscous Flow Within a 

Multistage Turbine,” Trans. ASME, 112, pp. 370–376.

2. Mielke M., Clark, D.S., and Wood, P., “RTA/GE57 Subscale Fan 

Rig, Final Report - Aero & Aeromechanical,” NASA/CR—2010-

215815.

3. Suder, K., Prahst, P., and Thorp, S., “Results of an Advanced Fan 

Stage Operating Over a Wide Range of Speed and Bypass Ratio—

Part I: Fan Stage Design and Experimental Results,” ASME, 

2010. GT2010-22825.

4. Shabbir, A., Zhu, J., and Celestina, M.L. “Assessment of the 

Three-Dimensional Models in a Compressor Rotor,” ASME, 

1996. ASME Paper No. 96–GT–198.

5. Shabbir, A. and Turner, M.G., “A Wall Function for Calculating 

the Skin Friction with Surface Roughness,” ASME, 2004. Paper 

GT2004–53908.

6. Spalding, D.B., “A Single Formula for the Law of the Wall,” 

1961, J. Appl Mech., pp. 455–458.

7. Kirtley, K.R., Beach, T.A., and Adamczyk, J.J. Numerical 

Analysis of Secondary Flow in a Two-Stage Turbine. AIAA, 

1990, AIAA–90–2356.

8. Shabbir, A., Adamczyk, J.J., Strazisar, A.J., and Celestina, M.L., 

“The Effect of Hub Leakage Flow on Two High-Speed Axial 

Flow Compressor Rotors,” ASME, 1997. Paper 97–GT–346.

9. VanZante, D., Strazisar, A., Wood, J., Hathaway, M., and Okiishi, 

T., “Recommendations for Achieving Accurate Numerical 

Simulation of Tip Clearance Flows in Transonic Compressor 

Rotors,” ASME, Oct. 2000, J. of Turbomachinery, Vol. 122, 

p. 733.

10. Adamczyk, J.J. Private Communication. 

11. Beach, T. Private Communication.

10

15

20

25

30

35

40

30% 40% 50% 60% 70% 80% 90% 100% 110%

Corr. Inlet MassFlow, kg/sec

Percent of Rotor Design Speed

DESIGN

DATA

APNASA

0%

10%

20%

30%

40%

50%

60%

70%

80%

90%

30% 40% 50% 60% 70% 80% 90% 100% 110%

Stall Margin

Percent of Rotor Design Speed

DESIGN

DATA

APNASA

0.3

0.4

0.5

0.6

0.7

0.8

0.9

30% 40% 50% 60% 70% 80% 90% 100% 110%

Adiabatic Efficiency

Percent of Rotor Design Speed

DESIGN

DATA

APNASA
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Results of an Advanced Fan Stage Operating Over a Wide Range of Speed and Bypass Ratio 

Part 2: Comparison of CFD and Experimental Results

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WBS 599489.02.07.03.07 

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14. ABSTRACT

NASA and GE teamed to design and build a 57 percent engine scaled fan stage for a Mach 4 variable cycle turbofan/ramjet engine for 

access to space with multipoint operations. This fan stage was tested in NASA’s transonic compressor facility. The objectives of this test 

were to assess the aerodynamic and aero mechanic performance and operability characteristics of the fan stage over the entire range of 

engine operation including: 1) sea level static take-off; 2) transition over large swings in fan bypass ratio; 3) transition from turbofan to 

ramjet; and 4) fan wind-milling operation at high Mach flight conditions. This paper will focus on an assessment of APNASA, a multistage 

turbomachinery analysis code developed by NASA, to predict the fan stage performance and operability over a wide range of speeds (37 to 

100 percent) and bypass ratios. 

15. SUBJECT TERMS

Fan stage; Compressor rotor 

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