Kenneth L. Suder

Glenn Research Center, Cleveland, Ohio

Patricia S. Prahst

AP Solutions, Inc., Cleveland, Ohio

Scott A. Thorp

Glenn Research Center, Cleveland, Ohio

Results of an Advanced Fan Stage Operating

Over a Wide Range of Speed and Bypass Ratio

Part 1: Fan Stage Design and Experimental Results

NASA/TM—2011-216769/PART1

October 2011

GT2010–22825
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Kenneth L. Suder

Glenn Research Center, Cleveland, Ohio

Patricia S. Prahst

AP Solutions, Inc., Cleveland, Ohio

Scott A. Thorp

Glenn Research Center, Cleveland, Ohio

Results of an Advanced Fan Stage Operating

Over a Wide Range of Speed and Bypass Ratio

Part 1: Fan Stage Design and Experimental Results

NASA/TM—2011-216769/PART1

October 2011

GT2010–22825

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

Authors would like to acknowledge General Electric (GE) personnel: Peter Wood and Mark Mielke for the many

helpful discussions related to the fan stage design and their insights into the interpretation of the experimental and 

computational results. Additional acknowledgement to Dr. Aamir Shabbir of United Technologies Research Center

(UTRC), Rick Mulac from the University of Toledo, Dr. John Adamczyk (NASA retired), and Tim Beach

 from AP Solutions for their support and efforts related to the APNASA fl ow simulations. Without these

 acknowledged contributions the work presented in this paper would not have been possible.

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

Trade names and trademarks are used in this report for identifi cation 

only. Their usage does not constitute an offi cial endorsement, 

either expressed or implied, by the National Aeronautics and 

Space Administration.

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—2011-216769/PART1 1 

Results of an Advanced Fan Stage Operating Over 

a Wide Range of Speed and Bypass Ratio 

Part 1: Fan Stage Design and Experimental Results 

Kenneth L. Suder 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Patricia S. Prahst 

AP Solutions, Inc. 

Cleveland, Ohio 44135 

Scott A. Thorp 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

NASA’s Fundamental Aeronautics Program is investigating 

turbine-based combined cycle (TBCC) propulsion systems for 

access to space because it provides the potential for aircraftlike, space-launch operations that may significantly reduce 

launch costs and improve safety. To this end, National 

Aeronautics and Space Administration (NASA) and General 

Electric (GE) teamed to design a Mach 4 variable cycle 

turbofan/ramjet engine for access to space. To enable the wide 

operating range of a Mach 4+ variable cycle turbofan ramjet 

required the development of a unique fan stage design capable 

of multi-point operation to accommodate variations in bypass 

ratio (10), fan speed (7), inlet mass flow (3.5), inlet 

pressure (8), and inlet temperature (3). In this paper, NASA 

has set out to characterize a TBCC engine fan stage 

aerodynamic performance and stability limits over a wide 

operating range including power-on and hypersonic-unique 

“windmill” operation. Herein, we will present the fan stage 

design, and the experimental test results of the fan stage 

operating from 15 to 100 percent corrected design speed. 

Whereas, in the companion paper (Ref. 1), we will provide an 

assessment of NASA’s APNASA code’s ability to predict the 

fan stage performance and operability over a wide range of 

speed and bypass ratio. 

Introduction 

Studies performed under NASA’s Next Generation Launch 

Technology Program and the NASP High Speed Propulsion 

Assessment (HiSPA) program indicated a variable cycle 

turbofan/ramjet was the best configuration to satisfy accessto-space mission requirements for the first stage of a twostage-to-orbit system because this configuration maximizes 

the engine thrust-to-weight ratio while minimizing frontal 

area (Ref. 2). To this end, NASA and GE teamed to design a 

variable cycle engine for an aircraft launch vehicle with 

Mach 4+ capability for access to space (Refs. 3 and 4). The 

flight envelope of a Mach 4+ like space launch vehicle 

operating from runway takeoff with continuous acceleration 

through transition from turbofan to ramjet operation requires 

a turbofan engine with wide operating range capability. To 

enable the wide operating range of a Mach 4+ variable cycle 

turbofan ramjet required the development of a unique fan 

stage design capable of multi-point operation which provided 

high pressure ratio and efficiency at takeoff through the mid 

range of engine operation, while avoiding stall and 

minimizing losses at the higher flight Mach numbers. 

To mitigate the risk of meeting the unique design 

requirements for the fan stage, NASA and GE teamed to 

design and build a 57 percent engine scaled fan stage to be 

tested in NASA’s transonic compressor facility. The goals of 

this test were to assess the aerodynamic and aero mechanic 

performance and operability characteristics of the fan stage 

over its required range of operation from 15 to 100 percent 

fan corrected speed. 

The objectives of this research activity was to assess and 

document the capability of state-of-the art design and analysis 

tools (validated for subsonic flight vehicles) to design and 

predict the performance and operability of an advanced fan 

stage designed to meet the requirements for the first stage of a 

two-stage-to-orbit hypersonic vehicle (i.e., necessitating wide 

multi-point operating range). These design and analysis tools 

are still relevant because the inlet has diffused the fan axial 

Mach number to subsonic, however the wide operating range 

and advanced configurations required for a Mach 4+ vehicle 

result in using these tools beyond the operating ranges these 

tools were validated. The ultimate goal being able to have 

confidence in the tools to design and analyze these advanced 

TBCC configurations to meet future mission requirements. 

In the following sections we will describe the fan stage 

design, the experimental facility and instrumentation, a 

sampling of test results to highlight the fan stage performance 
NASA/TM—2011-216769/PART1 2 

and operability characteristics, the pre-test CFD predictions 

and comparison to the experimental results, and conclude 

with a summary of lessons learned and rationale for 

additional post-test CFD simulations to assess the ability of 

our tools to predict the fan stage performance and operability. 

These post-test simulations will be presented in Part 2 of this 

paper (Ref. 1) and will include additional detailed 

comparisons of the simulations to the experimental results. 

Fan Stage Design 

An overview of the fan stage design requirements and 

traceability to a Mach 4 TBCC engine propulsion system is 

provided in References 3 to 6 where it is shown that the fan 

stage is a critical enabling component for the Mach 4 TBCC 

engine. The fan stage flow-path and components shown in 

Figure 1 are a 22 in. diameter scaled simulation of the engine 

(38 in. diameter) fan stage and includes the fan rotor, outlet 

guide vane (OGV), and splitter flow-path including the 

engine frame struts. To enable the wide operating range of the 

Mach 4+ capable engine required the development of a 

unique fan stage design for multi-point operation to 

accommodate variations in bypass ratio (10), fan rotor speed 

(7), inlet mass flow (3.5), inlet pressure (8), and inlet 

temperature (3). Herein, bypass ratio is defined as the ratio 

of mass flow in the bypass duct to that in the core duct 

(see Fig. 1).

These large variations of the inlet conditions and rotational 

speed introduced the following aerodynamic technical 

challenges to the fan stage design: 

1. Stall Free operation of the fan stage from 15 to 110 percent 

rotor design speed (corrected). 

2. Minimize the pressure losses through the fan rotor and 

OGV especially at the very high bypass ratios where the 

fan stage is at or approaching “windmill” conditions. 

3. Avoid choking and provide clean and stable flow in the 

bypass and core ducts throughout the 10 range of 

bypass ratio. 

4. Deliver the required inlet conditions to the downstream 

engine components (ramjet and core compressor), from 

takeoff through transonic and to ramjet operation. 

These technical challenges were augmented by the 

requirement to maintain traceability to the reference TBCC 

vehicle which resulted in 1) a fan stage design without inlet 

guide vanes; thereby, making it more difficult to maintain 

performance and operability over the wide operating range, 

and 2) a bypass duct arrangement that must maintain a 

diameter consistent with scale-up to the reference vehicle, yet 

not be so small as to incur huge pressure losses or potential 

flow choking resulting from high Mach numbers in the duct. 

Overall fan stage performance and operability therefore 

requires major consideration, as competing goals at different 

operating points become major drivers in the design. A 

summary of the fan stage design features is found in Table 1 

Figure 1.—Fan stage components and flow path. 

TABLE 1.—SELECTED FAN STAGE DESIGN PARAMETERS 

Fan rotor airfoils 25 

Tip speed at 17280 rpm 506 m/s (1660 ft/s) 

Radius ratio at LE 0.43 

Avg. aspect ratio 1.12 

Avg. solidity 2.2 

Tip clearance/chord ~0.009 nominal 

Specific flow 38.0 

OGV airfoils 48 

Design flow rate 37.3 kg/s (82 lbm/s) 

Stage PR 2.47 

Stage adiabatic efficiency 85.7 

Stall margin—high speed 12% 

Stall margin—low speed 20% 

and the resulting geometry for the fan and OGV are shown in 

Figure 2. The details of the mechanical design and 

aeromechanic test data are found in References 5 and 7, 

respectively. 

Fan Rotor Design 

The fan stage was designed by GE engineers following 

their standard design practice, with close collaboration from 

NASA engineers providing multi-stage CFD analysis in 

support of the design. The fan rotor was designed to produce 

a high pressure ratio (2.5) at lower flight Mach numbers while 

maintaining adequate stall margin (>10 percent) across a wide 

range of operating conditions. In order to deliver the required 

performance, an advanced technology, forward swept fan 

rotor design (Refs. 5 to 7) was employed. The resulting fan 

rotor geometry is shown in Figure 2(a). 

The fan stage design operating line (as determined by GE’s 

cycle code to meet mission requirements) is depicted by the 

black line connecting the black circles in Figure 3. CFD 

simulations were used to update the fan stage performance 

maps in the engine cycle deck. Single blade row CFD 

analyses were run using GE’s TACOMA (Ref. 8) code along 

the operating line at 100, 90, 80, 50, 37, 20, and 15 percent of 

design speed, as well as near stall at 100 and 80 percent speed. 
NASA/TM—2011-216769/PART1 3 

Figure 2.—Fan and OGV design airfoil shapes. 

Figure 3.—Fan stage operating map—design operating line.

Multi-stage CFD analyses were also conducted at select 

operating conditions (100, 50, and 37 percent of design 

speed), utilizing NASA’s APNASA (Ref. 9) code. These 

simulations predict that the fan stage should be capable of 

matching the design intent across the operating range. In 

order to insure adequate stall margin over the operating range, 

the fan stage test article provided the flexibility to incorporate 

various casing treatments over the fan rotor by installing a 

liner insert as depicted in Figure 1. Three liners were 

designed and tested with the fan stage: one smooth liner with 

a nominal fan tip clearance and 2 liners with identical 

circumferential grooves of which one had the same clearance 

as the smooth liner and the other with twice the clearance, 

herein designated open clearance. The 5 circumferential 

grooves over the rotor tip (see Fig. 1) have an axial width to 

radial height ratio of approximately 0.4 and are centered at 

10, 25, 40, 55, and 70 percent axial rotor chord. The 

sensitivity of stage performance and operability to liner 

configuration will be presented in the experimental results. 

OGV Design 

The OGV airfoil design is very three-dimensional, with 

significant amounts of bow and lean at the hub, both to 

strengthen the endwall flow and to help turn air into the 

engine core duct. The resulting geometry for the fan and 

OGV are shown in Figure 2. Along the operating line, shown 

in Figure 3, the mission requires the fan to operate at near 

peak efficiency at design speed (i.e., highly loaded operation), 

but near the maximum flow at 15 to 37 percent design speed 

Meridional

View Aft Looking

Forward

ALPHA

MUO 0

ETAO= 0

MU

ETA

Hub

Tip

LE TE

MUO= 0

ETAO= 0

UPSILON

S/L 13

mid-span

Hub Aft Looking

Forward

Meridional

View

Tip

a) Advanced technology forward swept fan b) OGV bowed and leaned at the hub to turn air 

into the core.

Meridional

View Aft Looking

Forward

ALPHA

MUO 0

ETAO= 0

MU

ETA

Hub

Tip

LE TE

Meridional

View Aft Looking

Forward

ALPHA

MUO 0

ETAO= 0

MU

ETA

Hub

Tip

LE TE

MUO= 0

ETAO= 0

UPSILON

S/L 13

mid-span

Hub Aft Looking

Forward

Meridional

View

Tip

MUO= 0

ETAO= 0

UPSILON

S/L 13

mid-span

Hub Aft Looking

Forward

Meridional

View

Tip

a) Advanced technology forward swept fan b) OGV bowed and leaned at the hub to turn air 

into the core.

% Corrected Fan Flow

0 20 40 60 80 100

3.00

2.50

2.00

1.50

1.00

.90

.50

Fan Pressure Ratio

• Windmill Region

PR < 1

• Design Fan PR = 2.45 for Base Map

- Single Stage

- No IGV

10% 20% 30% 40%

50% 60%

70%80% 90%

100 %Speed

Minimum Stall margin

or flutter margin range

Take-Off

Low bypass mode

BPR ~ 0.7

37% Speed Max 

Bypass Ratio ~7)

Potential Fan OGV and 

outer duct choking region

Maximum fan OGV

aero loading, ADP

Engine Cycle Predicted / Intent

Single Blade Row CFD Analysis (TACOMA)

Multistage Analysis (APNASA)

Adiabatic 

Efficiency 

Contours
NASA/TM—2011-216769/PART1 4 

(lightly loaded operation). Note that this operating line is 

unique to the hypersonic mission in that the fan bypass and 

core streams are independently throttled using a variable 

nozzle in each stream; thereby resulting in an operating line 

very much different from a fan operating line generated using 

a variable or fixed nozzle. Therefore, along the required fan 

operating line, the fan loading and (work factor) decreases 

with fan rotational speed which reduces the tendency of the 

fan rotor to stall at the lower speeds, but produces lower air 

angle onto the downstream OGV. At the aerodynamic design 

point (ADP), the OGV is heavily loaded and operates with 

high inlet air angles and Mach numbers, thereby resulting in 

high solidity to avoid the risk of stall. Quantitative values 

indicative of the OGV incidence and inlet Mach number will 

be presented in the experimental results session. 

The original design intent was to have a fixed OGV to 

minimize weight and complexity and minimize the clearance 

between the OGV and the endwalls. However, these large 

swings of incidence to the OGV over the 15 to 100 percent 

speed operating range, create significant aerodynamic design 

challenges resulting in the necessity for a variable OGV. To 

illustrate this point, the mid-span flow field of the OGV was 

analyzed using a two-dimensional CFD (2D-TACOMA) code 

(Ref. 8) at 37 percent design speed for the OGV setting angle 

of 0° (i.e., the baseline setting angle at the ADP) and a more 

open setting angle of 10°. These results, plotted in Figure 4, 

show that if the OGV was fixed at the 0° setting angle over 

the operating range, the high negative incidences onto the 

vane causes the pressure surface to separate. This in turn 

creates additional blockage through the vane creating a 

choking problem which would not only restrict flow to the 

downstream engine components and reduce thrust but also 

would impact the ability to move the fan operating point 

with downstream variable geometry. For the 10° open setting 

angle, these OGV flow field simulations, shown in Figure 4, 

illustrates the OGV still has a small but very limited flow 

separation off the pressure surface. To open the OGV further 

would reduce the flow separation at the expense of increasing 

endwall clearance and compromising design speed 

performance. 

Figure 5 compares the flow field solution (obtained from the 

three-dimensional N-S APNASA code (Ref. 9)) for the OGV 

design at 100 percent speed (0° setting angle) to the results at 

Figure 4.—Two-dimensional CFD (TACOMA) simulation of the 

OGV flow field at 50 percent span and 37 percent speed. 

Figure 5.—APNASA simulations of the OGV flow field at 100 and 37 percent speed for the fixed and open OGV settings. 

LE TE LE

OGV

5% Pitch

LE TE LE

OGV

50% Pitch

LE TE LE

OGV

50% Pitch 

LE TE LE

OGV

95% Pitch

LE TE LE

OGV

5% Pitch

LE TE LE

OGV

95% Pitch

Streamwise Direction

Radial Direction

37% Speed

100% Speed (ADP)

0.0 0.16 0.32 0.48 0.64 0.8

Normalized Axial Velocity Scale
NASA/TM—2011-216769/PART1 5 

37 percent speed with the OGV rotated 10° open. At 

100 percent design speed the OGV core flow field is well 

behaved, but the bypass duct flow field exhibits regions of 

low axial momentum in the bypass duct which results from 

the area increase and resulting diffusion in the bypass duct. 

Whereas, at 37 percent speed the opposite occurs in that the 

OGV flow field provides low axial momentum flow in the 

engine core duct, and the bypass duct flow is well behaved. 

The extremely low velocity in the lower spans of the OGV is 

required to obtain the high bypass ratio required for the cycle. 

Therefore, the final OGV design balances competing 

requirements over the entire range of operation by 

minimizing losses in the core duct at 100 percent speed where 

the bypass ratio is low and the engine thrust is derived mainly 

from the engine core, and it passes the flow without choking 

the bypass duct or incurring massive separations in the core 

duct where the bypass ratio is high and the thrust is derived 

mainly from the bypass flow. In summary, the final stage 

design satisfied its performance and operability requirements 

along the entire operating line. Through careful flow path and 

airfoil design, both single blade row and multistage CFD 

analyses showed that these requirements could be met. In the 

remainder of the paper, the experimental results will compare 

the measured performance to the design intent and show 

quantitatively the compromises in performance resulting from 

operating over such a wide operating range. 

Experimental Setup 

The NASA Single-Stage Axial Compressor Facility, 

illustrated in Figure 6, was used to investigate the 

aerodynamic performance of this compressor. The drive 

system consists of a 7,000 hp, 3,600 rpm electric drive motor 

and a 7,000 hp speed-increasing gearbox with a gear ratio of 

5.9, powered by the Engine Research Building (ERB) 

Variable Frequency system. The compressor speed can be 

controlled between 1,860 and 21,240 rpm. An in-line torque 

meter rated for 22,000 in.-lb measures power absorbed by the 

research compressor. The test rig shafting is bored to provide 

the capability for rotating measurements using a 100 channel 

slip ring. 

The facility air system is sized for a maximum flow of 

100 lbm/s. Atmospheric air is drawn into the facility, passes 

through a filter, through a flow measuring orifice plate, 

through inlet control valves, and into the plenum chamber. 

The plenum contains flow conditioning to provide an inlet 

flow turbulence intensity of 1.5 percent at the bell-mouth 

entrance and <1 percent at the fan rotor. The mass flow rate 

through the bypass and core ducts (refer to Fig. 1) is 

independently controlled by two throttle valves downstream 

of the fan stage. Air discharged from the exhaust collector is 

cooled by water spray injection and exhausted to atmosphere 

or via the altitude exhaust system. The altitude exhaust 

capability is a key feature of this facility which enabled 

mapping of fan stage operation at 50 percent design speed 

Figure 6.—NASA Single Stage Compressor Facility. 

and below where the fan stage pressure ratio is near or below 

1.0 (i.e., referred to as “windmill” condition in the context of 

this paper). 

Instrumentation 

Rig aerodynamic instrumentation was designed to capture 

flow details of interest and a schematic is presented in Figure 7. 

Full details of the instrumentation package can be found in 

References 5 and 10. In order to capture the inlet flow field, 

inlet rakes were placed at the bell-mouth exit. These rakes 

were placed well upstream of the fan to reduce the risk of 

potentially harmful aeromechanical interaction between these 

large rakes and the fan. High response pressure transducers 

and matching static pressure taps were placed in an axial 

array over the fan tip to allow mapping of the shock structure 

in this region, as a function of speed and throttle position. Six 

OGV leading edges were instrumented with radial arrays of 

total temperature and total pressure sensors to evaluate fan 

performance. Overall stage performance measurements were 

made with rakes located downstream of the OGV before the 

flow split (refer to Fig. 7). Eight rakes were clocked around 

the circumference with no more than one rake per OGV 

passage which could be reconstructed to measure across one 

OGV pitch (see the inset in Fig. 7). Each rake consisted of 7 

radial elements at 8, 24, 39, 53, 67, 81, and 94 percent spans 

(equal areas) measured both total pressure and temperature. 

Static pressures were measured at the hub and case location 

coinciding with the measurement plane of each rake and were 

used to quantify stage performance. The OGV’s were also 

instrumented with static pressure taps at 4 span-wise locations 

(10, 15, 25, and 45 percent span from the hub) from vane 

leading to trailing edge on both pressure and suction surfaces. 

These static taps were used to assess OGV blade loading as 

well as characterize the incidence angle on the OGV. 

Boundary layer and total temperature and pressure 

combination rakes were placed in the bypass and core ducts 

downstream of the splitter, both to determine duct and strut 
NASA/TM—2011-216769/PART1 6 

Figure 7.—Instrumentation schematic for fan stage used to obtain stage performance and operability. 

losses as well as to measure the mass flow rate in each stream 

so that the required operating bypass ratio could be set. 

Finally, static pressure taps were distributed along the flowpath in several critical regions to detect separated flow. 

Test Procedure 

Initial testing focused on defining the safe zones of 

operation and entailed mapping the operating line over the 

speed range. To limit the number of stalls, additional mapping 

from the operating line to the stall line were acquired while 

monitoring operability Kulites for aerodynamic stall and light 

probes, accelerometers, and strain gages for mechanical and 

aeromechanical stability. If any of these instruments indicated 

we were approaching and aerodynamic or aeromechanical 

instability the mapping was stopped and returned to the 

operating line. This procedure was used for the smooth liner 

and grooved liner with nominal clearance configurations, but 

for the grooved liner large clearance (2 times nominal) all but 

4 of the strain gages were inoperable and it was decided to 

only go to the same operating conditions that were cleared 

during the testing of the previous 2 liners. The aeromechanic 

results are documented in Reference 5 and show that the fan 

stage is safe to operate anywhere along the operating line. 

Furthermore, the aeromechanic results in Reference 5 show 

the fan stage can operate over the entire operating range 

without endwall casing treatment because the stage stalls 

aerodynamically prior to reaching aeromechanical stability 

limits (i.e., flutter boundaries). However, with casing 

treatment the aerodynamic stability limit and aeromechanical 

stability limit (stall flutter boundary) overlap and therefore 

limited near stall data were acquired with casing treatment 

configurations. In all cases the speed lines were mapped to 

very near aerodynamic stall conditions. 

In order to obtain data for a speed line, the pressure ratio, 

mass flow, and bypass ratio were set to match the design 

intent along the operating line. Initially, the intent was to map 

a speed line at constant bypass ratio. However, once testing 

began it was clear that manually adjusting the throttle valve 

for each of the bypass and core ducts would be too time 

consuming. Therefore, the speed lines were acquired by 

initiating at the operating line condition and then by first 

moving the bypass valve while maintaining a fixed position 

on the core duct throttle valve and second by returning to the 

operating line and closing the core duct valve while 

maintaining a fixed position on the bypass throttle valve. 

This process enabled full mapping from maximum flow to 

near stall condition at a given speed but does result in a slight 

variation of bypass ratio along the speed line. In all instances 

the maps at a given speed line will be presented with the data 

obtained with the bypass ratio closest to the bypass ratio of 

the design intent along the operating line. In the companion 

paper (Ref. 1), the sensitivity of the fan stage performance to 

slight variations in bypass ratio will be presented. 

Experimental Results 

Aerodynamic fan stage performance characteristics were 

acquired at 15, 25, 37, 50, 60, 70, 80, 85, 90, 95, 100, and 

101 percent of design rotor speed for three different fan rotor 

casing configurations corresponding to 1) smooth wall at 

nominal clearance, 2) circumferentially grooved casing at 

nominal clearance and 3) circumferentially grooved casing at 

twice nominal tip clearance. Nominal tip clearance was 

OGV Exit Pt & Tt

sensors 8 

Circumferential 

Locations

OGV Leading Edge 7 immersions

Pt & Tt sensors 3 

Circumferential 

Locations

7 immersions

Rotor

OGV

Over the Rotor Tip 

Statics and Kulites

5 circumferential 

grooves 

Aft looking upstream depicting reconstructed

Rake locations (as colored lines) covering one

OGV pitch

OGV TrailingEdges

Shroud

Hub Rake Locations
NASA/TM—2011-216769/PART1 7 

measured to be 0.005 of tip chord (design value = 0.009) at 

fan mid-chord at 100 percent speed; whereas, the open 

clearance was measured to be 0.018 of tip chord at 

100 percent speed. Uncertainty analysis was performed on the 

measured and relevant calculated parameters. The 

uncertainties actually vary for every measured data point; 

however, at a given rotational rotor speed the uncertainties 

remain nearly constant. The nominal data uncertainties for 

each speed are provided in Table 2. Only selected data are 

presented to highlight the features of the wide operating 

design requirements for this stage. For example, in the 

following sections we will discuss the impact of the liner 

configuration and OGV setting angle on the stage 

performance. Radial profiles of the total pressure ratio at 

100 percent speed are presented to indicate the distribution of 

OGV losses. Additional radial profiles will be presented in 

Part 2 of this paper to make detailed comparisons with the 

CFD simulations. 

TABLE 2.—NOMINAL DATA UNCERTAINTIES 

Measured Speed Lines and Comparison With 

Design Intent 

Experimental results of the fan stage characteristics at 15, 

25, 37, 50, 60, 70, 80, 85, 90, 95, and 100 percent of rotor 

design speed are shown in terms overall stage pressure ratio 

and adiabatic efficiency for the smooth wall configuration in 

Figure 8. The speed-lines are being shown for smooth wall 

nominal clearance (0.012 in. at 100 percent speed varying to 

0.026 in. at 50 percent speed and below). The black line in the 

pressure ratio plot represents the design operating line for the 

compressor and the red line indicates the notional 

aerodynamic stall line as estimated from GE’s design 

procedure. Note that the uncertainties in pressure ratio and 

mass flow are within the symbol size. The uncertainty in 

efficiency varies dramatically with speed (Table 2) primarily 

because the work done on the fluid diminishes rapidly with 

rotational speed and the uncertainty in temperature 

measurement results in greater uncertainties in efficiency at 

the lower speeds. Also note the efficiency at 15 percent speed 

is not shown because at this speed the stage pressure ratio is 

less than 1 and stage losses are greater than the work added to 

the fluid by the rotor and therefore the meaning of efficiency 

(ratio of energy added to the fluid to the energy input to drive 

the fan rotor) is invalid. For simplicity the data presented 

hereafter, will not include the error bars. The overall 

performance characteristics of the fan stage for the grooved 

liner with nominal and open clearances are provided in 

Figures 9 and 10. 

The data point corresponding to the lowest value of 

corrected flow for each speed-line does not necessarily 

correspond to the aerodynamic stall point. If an aerodynamic 

stall was obtained, it is shown by the solid red filled symbol 

data point as indicated in the pressure ratio plot for the speedlines of 95, 90, 85, 80, 70, 50, and 37 percent of fan rotational 

speed. As can be seen from Figure 8, 100 percent speed was 

not throttled to an aerodynamic stall, instead an audible hum 

was observed. Data from the strain gages, light probes, and 

Kulite data indicated the hum was not associated with an 

increase in fan stage mechanical stress nor of that for a typical 

rotating stall. However the operability Kulite measurements 

exhibited an increase in unsteadiness at the last data point 

acquired at 100 percent speed in Figure 8 for the smooth wall 

configuration consistent with those observed at a near stall 

operating condition. 

Detailed comparison of measured performance along the 

operating line from 15 to 100 percent speed is shown in 

Figure 11. It is evident the stage design met its design intent 

not only in pressure rise and efficiency but also in flow rate 

and stall margin. 

Figure 8.—Stage operating characteristics for the smooth 

nominal clearance liner configuration. 

DATA Uncertainties

% Rotational Speed 100% 90% 80% 60% 37%

MassFlow (kg/sec) 0.15 0.16 0.19 0.23 0.03

PR 0.008 0.008 0.008 0.008 0.008

Efficiency 0.005 0.006 0.008 0.014 0.03

Stall MassFlow (kg/sec) 0.15 0.16 0.19 0.23 0.03

Stall PR 0.008 0.008 0.008 0.008 0.008

StallMargin (%) 0.40% 0.50% 0.70% 1.10% 0.20%

0.6

0.65

0.7

0.75

0.8

0.85

0.9

0 5 10 15 20 25 30 35 40

Adiabatic Efficiency

Corrected Inlet Flow, kg/sec

RTA Fan Smooth Casing Adiabatic Efficiency

100%

90%

80%

85% 95%

50% 60%

70%

37%

0.5

1

1.5

2

2.5

3

0 5 10 15 20 25 30 35 40

Total Pressure Ratio

Corrected Inlet Flow, kg/sec

RTA Fan Smooth Casing Stage Pressure Ratio

90%

80%

85%

95%

50%

60%

70%

25%

37%

15%

100%

90%

80%

85%

95%

50%

60%

70%

25%

37%

15%

Operating Line – design 

Stall line – design est.

Measured Stall 

OGV at 0 degrees

OGV at 10 degrees open
NASA/TM—2011-216769/PART1 8 

Figure 9.—Stage operating characteristics for the 

grooved nominal clearance liner configuration. 

Figure 10.—Stage operating characteristics for the 

grooved open clearance liner configuration. 

Figure 11.—Comparison of design intent to the 

measured performance and operability parameters 

along the operating line for the smooth wall liner 

configuration. 

 

Operating Line – design 

Stall line – design est.

Measured Stall 

OGV at 0 degrees

OGV at 10 degrees open

0.5

1

1.5

2

2.5

3

0 5 10 15 20 25 30 35 40

Total Pressure Ratio

Corrected Inlet Flow, kg/sec

RTA Fan Grooved Casing Stage Pressure Ratio 100%

90%

80%

85%

95%

50%

60%

70%

25% 37% 15%

0.6

0.65

0.7

0.75

0.8

0.85

0.9

0 5 10 15 20 25 30 35 40

Adiabatic Efficiency

Corrected Inlet Flow, kg/sec

RTA Fan Grooved Adiabatic Efficiency

60%

37%

50%

70%

80%

85%

90%

95%

100%

0.6

0.65

0.7

0.75

0.8

0.85

0.9

0 5 10 15 20 25 30 35 40

Adiabatic Efficiency

Corrected Inlet Flow, kg/sec

RTA Fan Grooved Casing 

Open Clearance Adiabatic Efficiency

60% 70%

100%

85%

90% 95%

0.5

1

1.5

2

2.5

3

0 5 10 15 20 25 30 35 40

Total Pressure Ratio

Corrected Inlet Flow, kg/sec

RTA Fan Grooved Casing - Open Clearance 

Stage Pressure Ratio

100%

95% 90%

85%

70%

60%

37% 50%

15% 25%

Operating Line – design 

Stall line – design est.

Measured Stall 

OGV at 0 degrees

OGV at 10 degrees open

0.3

0.4

0.5

0.6

0.7

0.8

0.9

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

Efficiency vs. Speed at Trajectory

DESIGN

DATA

20

30

40

50

60

70

80

90

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

Corr. MassFlow vs. Speed at Stall

DESIGN

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%

PR vs. Speed at Trajectory

DESIGN

DATA

0.00%

10.00%

20.00%

30.00%

40.00%

50.00%

60.00%

70.00%

80.00%

90.00%

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

Stall Margin vs. Speed

DESIGN

DATA

Corrected Speed
NASA/TM—2011-216769/PART1 9 

Sensitivity of Performance to Outer Endwall 

Configuration 

Based on the experimental results presented in Figures 8 to 

11 it is evident that the fan stage meets or exceeds the 

operating line requirements over the entire range of 15 to 

100 percent fan speed. (In the companion paper (Ref. 1) of 

this paper the design intent, experimental data, and APNASA 

results explicitly show this agreement at 100, 80, and 

37 percent speed.) It is also evident that the stall points for the 

grooved liner configurations occur at a lower operating mass 

flow than the smooth wall configuration as expected for 

circumferential groove casing treatment. Furthermore, the 

maximum flow achieved by the grooved liners configurations 

is about 1 percent less than the smooth wall configuration for 

the 70 to 100 percent design speed lines. The sensitivity of 

the fan stage performance to the 3 liner configurations is 

more accurately quantified from the plots of stage pressure 

ratio and adiabatic efficiency at 90 percent speed presented in 

Figure 12. Note the last data point is at the near stall point and 

the data missing at the higher flow rate is consistent with the 

test procedure. The maximum adiabatic efficiency is 2 points 

higher for the smooth wall configuration at the higher flow 

rate near the operating line consistent with the higher pressure 

ratio. However, from the operating line towards the stall line 

the efficiency is nearly the same for the grooved and smooth 

wall liners. Though not shown explicitly the same results as 

discussed above for 90 percent speed were observed at 

95 percent speed. Figure 13 shows the measured casing static 

pressure distribution at 90 percent speed (normalized by the 

inlet total pressure) over the fan tip for the three liner 

configurations. The casing statics of Figure 13 are consistent 

with the overall averaged total pressure rise characteristics 

shown in Figure 12. The static pressure rise is indicative of 

fan shock location such that at choke conditions the shock is 

oblique and pulled back in the fan passage such that the shock 

hits the suction surface near the trailing edge; whereas at near 

stall the shock is more normal and intersects the suction 

surface further upstream nearer to the leading edge. The 

influence of the circumferential grooves on the stability of the 

fan stage is evidenced by comparing the results in Figures 12 

and 13. On the high flow rate side of the characteristic the 

grooves and open clearance act as a blockage and reduce the 

choking flow rate as evidenced by the same fan exit pressure 

at a much lower flow rate for the grooved liners. In other 

words the effect of the grooves at higher flow rates requires 

the fan to operate at a lower flow (higher incidence) to obtain 

the same fan exit pressure. Comparing results at near stall 

conditions, the shock for the grooved liner with open 

clearance is forward of the rotor leading edge, yet the fan 

stage is not stalled; whereas for the smooth liner 

configuration the shock is well within the passage and on the 

verge of stall at a higher flow rate relative to the grooved liner 

near stall flow rate. Therefore, to summarize, the relief 

provided by the grooves to enable lower flow rates and more 

Figure 12.—Comparison of stage pressure ratio and adiabatic 

efficiency at 90 percent speed for the 3 liner configurations. 

Figure 13.—Comparison of normalized fan tip static pressure 

distribution for the 3 liners tested at 90 percent speed. 

 

1.7

1.8

1.9

2

2.1

2.2

2.3

25 30 35 40

Total Pressure Ratio

Corrected Inlet Massflow, kg/sec.

Grooved Liner (506) Smooth Liner (508) Grooved Liner open cl (507)

0.6

0.65

0.7

0.75

0.8

0.85

0.9

25 30 35 40

Adiabatic Efficiency

CorrectedMassflow, kg/sec.

Grooved Liner (506) Smooth Liner (508) Grooved Liner open cl (507)

90% Speed

90% Speed

0.6

0.8

1

1.2

1.4

1.6

1.8

2

‐20 ‐10 0 10 20 30 40 50 60 70 80 90 100

Ps/Pt (Over Rotor)

% Rotor Chord

Grooved, Choke flow Grooved,Operating Line

Grooved, Near Stall Smooth, Choke Flow

Smooth,Operating Line Smooth, Near Stall

Grooved, open clearance, Choke Flow Grooved,Open Clearance, Operating  Line

Grooved,Open Clearance, Near Stall
NASA/TM—2011-216769/PART1 10 

stability margin at near stall operation is offset by increased 

blockage (resulting in a reduction of the maximum flow) and 

a lower efficiency at higher flow rates. 

The mass flow, pressure ratio, and efficiency along the 

operating line are tabulated and compared to the 

corresponding conditions at stall for each of the liner 

configurations in Table 3. The stall margin is calculated with 

the operating line condition selected as the reference 

condition using the standard NASA definition of the stall 

margin: 















   1

PR @ref

Massflow@ref

Mass flow@stall

PR @stall Stallmargin

TABLE 3.—SUMMARY OF TIP CLEARANCE AND 

ENDWALL TREATMENT ON STAGE PERFORMANCE 

AND OPERABILITY 

Recall from the section describing the test procedure that 

the fan stage was only throttled to stall for a select few 

operating conditions. From Table 3 it is apparent that the 

grooved liners increase the stall margin by 5 to 6 percent at 

80 to 100 percent speed, which is typical for circumferential 

groove casing treatments. For the grooved liner 

configurations at 100 percent speed, where the clearance 

varied from 0.012 to 0.045 in., the stall margin and efficiency 

was nearly identical; thereby, indicating that at the higher 

loaded conditions, there is little sensitivity to clearance. (Note 

that aside from the fan tip clearance the two grooved liner 

configurations were identical.) Below 70 percent fan speed 

there was virtually no difference measured in fan stage 

performance and operability for the 3 different liner 

configurations. 

Sensitivity of OGV to Stage Performance and 

Operability 

Static pressure taps on both the suction and pressure 

surface of the OGV were located at 10, 15, 25, and 45 percent 

span from the hub of the OGV to monitor the loading of the 

OGV in the core duct. To assess the loading and incidence as 

well as determine if the duct was choked the static taps were 

used to calculate an isentropic relative Mach number using 

the total pressure at the OGV leading edge. The results for the 

37, 60, 80, 90, and 100 percent speed operation are depicted 

in Figure 14. Note the OGV setting angle was open to 10° at 

37 percent speed condition to avoid the OGV leading edge 

separation and resulting blockage as predicted in the design as 

seen in Figure 4. At 37 and 60 percent speed along the 

operating line the stagnation streamline falls well on the 

suction surface of the OGV possibly resulting in acceleration 

around the leading edge to the pressure surface and potential 

separation and increased blockage as shown in Figures 4 and 

5. Note that at 100 percent speed the Mach number is nearly 

one along the suction surface and is indicative of higher OGV 

losses. Similarly, at 90 percent speed the Mach number is 

above 0.8 over much of the suction surface. Though not 

shown at 10 percent span and 100 percent speed the suction 

surface Mach number is above Mach 0.8 over 75 percent of 

the OGV chord and exceeds Mach 1 over the first 20 percent 

of the OGV chord length. Therefore at design speed the OGV 

passage is potentially choked in the lower spans. In 

conclusion these Mach number distributions are consistent 

with the design-predicted trends in OGV incidence and 

loading from 37 to 100 percent speed. 

To further show the sensitivity to OGV incidence and 

Mach number and clearly demonstrate the necessity for a 

variable OGV, the Mach number distribution at 60 percent 

speed for the OGV at 0°, 5°, and 10° is presented in 

Figure 15. At the design setting of 0° the flow is incident on 

the suction side and the Mach number approaches 1; whereas 

at 10° open the flow is better aligned with the OGV leading 

edge and the Mach number over the suction surface is much 

reduced. At the lower spans the results (not shown) show a 

much larger mismatch at the OGV leading edge indicating the 

OGV needs to swing more than 10° to obtain a near zero 

incidence along the operating line. With regards to the overall 

performance, the total pressure ratio maps at 50 and 

37 percent speed are plotted in Figure 16 for the baseline 

OGV setting angle of 0° as well for 5° and 10° open. These 

results, shown in Figure 16, show an increase in flow rate and 

increase in stage pressure rise for a given flow rate over most 

of the operating range. As already mentioned there is no 

effect of the liner configuration on stage pressure rise below 

70 percent speed. In summary these results indicate there is a 

need for a variable OGV to match the wide swings in 

incidence and inlet Mach number over the wide operating 

range and the OGV (not the fan) is the airfoil most sensitive 

to variations in stage performance over this operating range. 

grooved nominal clearance = 0.012" at 100%speed

DATA 100% 95% 90% 80% 50% 37%

Flow (lbm/sec) 82.6 78.77 72.16 63.1 46.03 39.2

PR 2.432 2.35 2.22 1.859 1.157 1.015

Efficiency 0.793 0.85 0.848 0.848 0.762 0.199

Stall Flow 78.26 73.05 61.6 52.1 27.7 19.74

Stall PR 2.68 2.44 2.222 1.88 1.264 1.138

StallMargin (%) 17.1 22 

Smooth Wall ‐ nominal Clearance = 0.012" at 100% speed

DATA 100% 90% 80% 50% 37%

Flow (lbm/sec) 82.5 71.84 62.24 45.63 38.3

PR 2.41 2.2 1.855 1.17 1.03

Efficiency 0.828 0.849 0.857 0.815 0.372

Stall Flow 79.14 64 53.9 30.72 22.3

Stall PR 2.6 2.2 1.87 1.264 1.139

StallMargin (%) 12.6 15.5

Grooved open clearance = 0.045" at 100% speed

DATA 101% 95% 90% 50% 37%

Flow (lbm/sec) 82.96 78.04 71.9 43.16 35.8

PR 2.442 2.335 2.167 1.133 1.003

Efficiency 0.812 0.852 0.851 0.599 0.039

Stall Flow 80.87 67.4 61.5 27.38 21.01

Stall PR 2.64 2.376 2.165 1.27 1.144

StallMargin (%) 16.9

Not Stalled

80% speed 

Stall Margin

90% Speed 

Stall Margin
NASA/TM—2011-216769/PART1 11 

Figure 14.—OGV surface isentropic Mach 

number values for the smooth wall nominal 

clearance liner configuration at 45 percent 

span for 37 to 100 percent speed. 

Figure 15.—Effect of OGV setting angle on isentropic Mach 

number distribution at 60 percent speed. 

Figure 16.—Sensitivity to stage pressure ratio and mass flow 

to OGV setting angle at 37 and 50 percent speed. 

0.0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

1.6

1.8

0 20 40 60 80 100

Mach Number

% Chord

37% Speed,

45% OGV Span

SS, Max Flow PS, Max Flow

SS, Op Line PS, Op Line

SS, Near Stall PS, Near Stall

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

0 20 40 60 80 100

Mach Number

% Chord

60% Speed,

45% OGV Span

SS, Max Flow PS, Max Flow

SS, Op Line PS, Op Line

SS, Near Stall PS, Near Stall

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

0 20 40 60 80 100

Mach Number

% Chord

80% Speed, 

45% OGV Span

SS, Max Flow PS, Max Flow

SS, Op Line PS, Op Line

SS, Near Stall PS, Near Stall

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

1.8

0 20 40 60 80 100

Mach Number

% Chord

90% Speed,

45% OGV Span

SS, Max Flow PS, Max Flow

SS, Op Line PS, Op Line

SS, Near Stall PS, Near Stall

0

0.2

0.4

0.6

0.8

1

1.2

1.4

1.6

0 20 40 60 80 100

Mach Number

% Chord

100% Speed, 

45% OGV Span

SS, Max Flow PS, Max Flow

SS, Op Line PS, Op Line

SS, Near Stall PS, Near Stall

0

0.2

0.4

0.6

0.8

1

1.2

0 20 40 60 80 100

Mach Number

% Chord

60% Speed, OGV = 0

45% span,suction 45% span, pressure

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

0.9

0 20 40 60 80 100

Mach Number

% Chord

60% Speed, OGV = 5 

45% span, suction 45% span, pressure

0

0.2

0.4

0.6

0.8

0 20 40 60 80 100

Mach Number

% Chord

60% Speed, OGV = 10 

45% span, suction 45% span, pressure

0.80

0.85

0.90

0.95

1.00

1.05

1.10

1.15

1.20

1.25

1.30

8.0 10.0 12.0 14.0 16.0 18.0 20.0 22.0 24.0

Total Pressure Ratio

Corrected Inlet flow, kg/sec

Grooved, 0

Grooved, -5

Grooved, -10

Smooth, 0

Smooth, -5

Smooth, -10

Grooved, open clearance, 0

Grooved, 0

Grooved, -5

Grooved, -10

Smooth, 0

Smooth, -5

Smooth, -10

Grooved, open clearance, 0

37%

Speedlines

50% 

Speedlines
NASA/TM—2011-216769/PART1 12 

The stage performance is very sensitive to OGV setting 

angle as depicted in Figures 14 to 16. Herein we will look at 

the results for 100 percent speed. Radial distributions of total 

pressure ratio at the OGV leading edge and downstream rake 

location (see Fig. 7) as the rotor is back-pressured from choke 

to stall is plotted in Figure 17. These results indicate as the 

stage is throttled nearly all of the increase in pressure occurs 

in the outer spans (beyond point B in the map). In addition 

there is evidence that the flow is actually choked in that the 

pressure ratio at about 40 percent immersion does not change 

with backpressure (also indicated by OGV surface statics). 

Note the total loss coefficient is defined as the difference in 

total pressure across the OGV divided by the OGV inlet 

dynamic pressure. Most surprising is the level of loss 

coefficient across the entire span greatly exceeding typical 

values of 0.02 to 0.2 found in traditional OGV designs. These 

higher losses are due in part to the higher inlet Mach numbers 

(refer to Fig. 14) which result in higher losses in the OGV 

across the span and not just at the endwalls. The losses are 

also in part due to the inherent compromise in performance to 

provide operability over such a large range but also reflect the 

slight mismatch between the fan and OGV. Any separations 

in the OGV are exacerbated by the rapid diffusion 

downstream of the OGV in the flow path (see Fig. 7) which 

was necessary to achieve the large swings in bypass ratio 

required by the cycle deck to meet mission requirements. 

These results are consistent with the sensitivity of the OGV to 

separation as it undergoes large swings in incidence and 

loading. Note that the loss coefficient plot was truncated at 

0.5 to show the distribution of loss near the operating line and 

towards stall. The maximum total loss coefficient reached 1.4 

in the hub for the wide open throttle setting and was due to 

massive separations in the lower spans of the OGV. Though 

wide open throttle is not an important operating condition, the 

loss variation in Figure 17 substantiates the sensitivity of 

stage performance to the OGV inlet conditions. In summary, 

these results show the sensitivity of the stage performance to 

OGV inlet conditions and indicate there is a slight mismatch 

between the OGV and fan over the wide operating range. The 

pre-test CFD is further investigated to shed insight into the 

reasons for this mismatch and exceedingly high losses in the 

OGV especially at 100 percent speed. 

Comparison of Data to Pre-Test CFD Using 

APNASA 

As was shown in the section on the fan stage design both 

NASA and GE CFD codes were used to assess the design as 

well as to provide input to the design system compressor 

maps. In this section we will discuss these pre-test CFD 

simulations using the NASA code APNASA. APNASA 

(Ref. 9) is a multistage analysis code which provided the 

ability to analyze the complete fan stage consisting of the fan 

rotor, OGV, and splitter/strut mid-frame assembly (see Fig. 1). 

Figure 17.—Radial distributions of total 

pressure ratio and loss coefficient 

across the OGV at 100 percent speed. 

A

B

C

D

E G F

1.5

1.7

1.9

2.1

2.3

2.5

2.7

2.9

35 35.5 36 36.5 37 37.5 38

Total Pressure Ratio

Corrected Inlet Mass Flow, kg/sec

100% Speedline

0

10

20

30

40

50

60

70

80

90

100

2 2.2 2.4 2.6 2.8 3 3.2 3.4

% Span from Hub

Total Pressure Ratio

100% Speed OGV Leading Edge 

A

B

C

D

E

F

G

0

10

20

30

40

50

60

70

80

90

100

1.4 1.6 1.8 2 2.2 2.4 2.6 2.8

% Span from Hub

Total Pressure Ratio

100% Speed OGV Exit Plane

A

B

C

D

E

F

G

0

10

20

30

40

50

60

70

80

90

100

0 0.1 0.2 0.3 0.4 0.5

% Span from Hub

OGV Loss Coefficient

Loss Coefficient 

A

B

C

D

E

F

G
NASA/TM—2011-216769/PART1 13 

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 the entire flow field. The APNASA fan 

stage simulation 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. 9). The standard twoequation – model was used to calculate the turbulent eddy 

viscosity coefficient as described by Shabbir et al. (Refs. 11 

and 12). Spalding’s formula (Ref. 13) 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, 

radial and tangential flow angle which were provided by GE 

from their design code. The flow in the clearance gap was 

simulated using a model suggested by Kirtley et al. (Ref. 14). 

APNASA results for the overall stage pressure ratio and 

adiabatic efficiency are compared to the experimental results 

of the fan stage characteristics at 80, 85, 90, 95, and 

100 percent of rotor design speed in Figure 18. The solid 

symbols represent the data and the open symbols connected 

by a line indicate the CFD results. Note that only at 100 and 

80 percent speed was the CFD simulations back-pressured 

until they reached their stability limit indicative of an 

aerodynamic stall point. In general the CFD compared 

favorably with the data along the operating line both in terms 

of mass flow, pressure ratio, and to a lesser degree with the 

efficiency. The major concern for the CFD was the inability 

of the CFD to predict the stall margin at 100 percent design 

speed; whereas at 80 percent speed the stall margin was well 

predicted. 

Additional analysis of the CFD at 100 percent speed 

indicated that if the rotor was back-pressured beyond the 

operating point at 100 percent speed, the flow in the lower 

50 percent span in the OGV became highly separated (due to 

the higher incidence) and as it was further diffused, the low 

momentum region grew rapidly and would not allow the CFD 

simulation to converge. These analyses are presented in 

Figure 19. CFD simulations were run on the fan rotor only 

and since the OGV was not in the flow-path the separation 

did not occur and the fan rotor could be back-pressured such 

that the stall margin was comparable to that of the measured 

stage data. These fan only simulations provided confirmation 

that the separation in the OGV, as predicted by the CFD for 

the stage configuration, was the reason the operating range at 

100 percent speed did not compare well with the data. Due to 

the sensitivity of the OGV loading and losses with incidence 

angle and backpressure (discussed in detail in the 

Experimental Results section), it was recognized that a mass 

Figure 18.—Comparison of pre-test CFD predictions 

compared to measured data at 80 to 100 percent 

speed. 

Figure 19.—Pre-test APNASA CFD results at 100 percent 

speed obtained by setting downstream pressure in the core 

and bypass ducts. 

 

1.5

1.7

1.9

2.1

2.3

2.5

2.7

0.6 0.65 0.7 0.75 0.8 0.85 0.9 0.95 1 1.05

Corrected Inlet Massflow, lbm/sec

Total Pressure Ratio RTA Fan

80%  

90%

95%  

100%  

Solid = DATA

Open = CFD

2.7

2.5

2.3

2.1

1.9

1.7

1.5

TOTAL PRESSURE RATIO 0.6

0.65

0.7

0.75

0.8

0.85

0.9

0.95

0.6 0.65 0.7 0.75 0.8 0.85 0.9 0.95 1 1.05

Corrected Inlet Massflow, lbm/sec

Adiabatic Efficiency

ADP

Normalized Inlet Mass Flow

Note:  

CFD only  taken

to Stall at 80% speed

80%  

90% 95%  

100%  

Solid = DATA

Open = CFD

0.95

0.90

0.80

0.75

0.70

0.65

0.85

0.60 0.70 0.80 0.90 1.00

0.60

Corrected  Flow Normalized Mass Flow

Adiabatic Efficiency

SS

SS

a) Last Converged Point

b) Further Back Pressuring

5% Span Behind TE of OGV
NASA/TM—2011-216769/PART1 14 

flow boundary condition must be implemented in the code to 

enable the CFD simulation to obtain the mass flow and 

bypass ratio demonstrated in the experimental data obtained 

away from the operating line. Further analysis of the CFD 

simulations as well as comparisons with the experimental 

results led to the conclusions that the CFD did not utilize the 

same blade geometry or boundary conditions consistent with 

the experimental data. Therefore, in the companion paper 

(Ref. 1), we will provide a detailed description of these 

discrepancies and will perform post-test CFD to assess 

NASA’s APNASA code’s ability to predict the fan stage 

performance and operability over a wide range of speed and 

bypass ratio. 

Summary and Concluding Remarks 

The fan stage from a state of the art Mach 4 turbine engine 

was scaled (57.3 percent linear scale), fabricated, and tested 

in the NASA W8 Single Stage Compressor facility. The fan 

stage was designed for multi-point operation to accommodate 

variations in bypass ratio (10), fan rotor speed (7), inlet 

mass flow (3.5), inlet pressure (8), and inlet temperature 

(3). Aerodynamic fan stage performance characteristics at 

15, 25, 37, 50, 60, 70, 80, 85, 90, 95, and 100 percent of rotor 

design speed, encompassing a bypass ratio swing from 0.7 to 

7, were acquired. These experimental results of the fan stage 

characteristics were acquired for three different fan rotor 

casing configurations corresponding to 1) smooth wall at 

nominal fan tip clearance, 2) circumferentially grooved 

casing at nominal tip clearance, and 3) circumferentially 

grooved casing with a more open clearance. The liner 

influenced the stage performance and operability for fan 

speeds greater than 70 percent speed, but no influence was 

measureable below 70 percent speed. At 90 to 100 percent 

speeds, the relief provided by the grooves to enable lower 

flow rates and more stability margin at near stall operation 

was offset by increased blockage (resulting in a reduction of 

the maximum flow) and a lower efficiency at higher flow 

rates. The experimental data agreed favorably with the design 

intent in terms of pressure ratio, efficiency and mass flow 

along the operating line. It was shown that the OGV loading 

and losses were sensitive to OGV setting angle and that a 

variable OGV was required to match the swings in incidence 

and inlet Mach number over the wide operating range. 

Furthermore, it was shown that the OGV (not the fan) is the 

airfoil most sensitive to variations in stage performance over 

this wide operating range. 

Pre-test CFD simulations were integral to the design effort 

and the results compared very well to the experiment along 

the design operating line. However, there was some 

discrepancy between the data and CFD in pressure ratio and 

adiabatic efficiency at off-design conditions. Most alarming 

was the fact that the pre-test CFD did not predict the stall 

margin at 100 percent speed. Inspection of the hardware and 

analysis of the data revealed the pre-test CFD did not have the 

corrected hot geometry blade shape and was run at a much 

larger fan tip clearance than the experiment. Furthermore, the 

experimental data and pre-test CFD results were obtained at 

different bypass ratios along the speed-lines for all data points 

other than for the bypass ratio that was specified in the design 

along the operating line for each fan speed. These 

discrepancies between the boundary conditions and input 

parameters measured in the experiment versus those used for 

the pre-test CFD predictions, necessitated additional post-test 

CFD simulations for detailed comparisons between the CFD 

simulations and experiment which will be presented in Part 2 

of this paper (Ref. 1). 

The future work will include additional testing and analysis 

to assess sensitivity of performance and operability to both 

endwall variations and inlet flow distortions. The distortion 

measured from the TBBC large scale inlet described in 

References 15 to 17 will be simulated with distortion screens 

at the inlet of the fan stage. The data will be used to evaluate 

the impact of distortion on the fan stage performance and 

operability as well as be used to assess the ability of the SOA 

tools to predict these effects. For non-axisymmetric inlet 

distortions unsteady analysis codes will need to be used. 

Therefore, studies have begun to baseline a full unsteady 

simulation by comparing the predictions of the unsteady 

codes (for the uniform inlet flow data presented herein) to 

those of APNASA’s averaged passage results presented in the 

companion paper (Ref. 1) to this paper. 

References 

1. Celestina, M.L., Suder, K.L., and Kulkarni, S. 2010, 

“Results of an Advanced Fan Stage Operating Over a 

Wide Range of Speed & Bypass Ratio—Part 2: 

Comparison of CFD and Experimental Results,” ASME 

Paper No. GT2010–23386, 2010. 

2. Bilardo V.J., Curran F.M., Lovell N.T., and Maggio G., 

2003, “Benefits of Airbreathing Launch Systems for 

Access to Space,” AIAA–2003–5265, 39th Joint 

Propulsion Conference in Huntsville, Alabama, July 20–

23, 2003. 

3. Shafer D.G., McNelis N., 2003, “Development of a 

Ground Based Mach 4+ Revolutionary Turbine 

Accelerator Technology Demonstrator (RTATD) for 

Access to Space,” ISOABE 2003. 

4. “RTA Propulsion System Data Requirements/Boeing 

TSTO Study from NASA Glenn,” RASER Contract 

NAS3-01135, 2002. 

5. Mielke, M., Clark, D., and Wood, P. 2009 “RTA/GE57 

Subscale Fan Rig, Final Report—Aero & 

Aeromechanical,” NASA/CR—2010-215815. 

6. Suder, K.L., Adamczyk, J.A, Wood, p., 2005 “Design and 

Analysis of a Mach 4 Capable Fan Stage for the RTA 

Turbine Based Combined Cycle Propulsion System,” 

JANNAF (Joint Army, Navy, NASA, Air Force) 42nd 
NASA/TM—2011-216769/PART1 15 

CS/30th APS/24th PSHS Joint Meeting, Charleston, SC 

June, 2005. 

7. Mielke, M., 2009 “RTA/GE57 Subscale Fan Rig, Final 

Report—Mechanical Design,” NASA/CR—2010-

215814. 

8. Smith, L.H. and Yeh, H., 1963, “Sweep and Dihedral 

Effects in Axial-Flow Turbomachinery,” ASME Journal 

of Basic Engineering, Vol. 85, pp. 401–416. 

9. Adamczyk, J.J., Celestina, M.L., Beach, T.A., and 

Barnett, M., 1990,” Simulation of Three-dimensional 

Viscous Flow within a Multistage Turbine,” Tran. ASME, 

112, 370–376. 

10. Mielke, M., Clark, D., and Wood, P. 2009 “RTA/GE57 

Subscale Fan Rig, Final Report—Aerodynamic Test 

Results,” NASA/CR—2009-215816. 

11. Shabbir, A. and Turner, M.G., 2004, “A Wall Function for 

Calculating the Skin Friction With Surface Roughness,” 

A. Shabbir and M.G. Turner, ASME Paper GT2004–

53908. 

12. Shabbir, A., Zhu, J., and Celestina, M.L. 1996, 

“Assessment of the Three-Dimensional Models in a 

Compressor Rotor,” ASME Paper No. 96-GT-1998. 

13. Spalding, D.B., 1961, “A Single Formula for the Law of 

the Wall,” J. Appl Mech. 28, pp. 455–458. 

14. Kirtley, K.R., Beach, T.A., and Adamczyk, J.J., 1990, 

“Numerical Analysis of Secondary Flow in a Two-Stage 

Turbine,” Paper no. AIAA–90–2356. 

15. Sanders, B.W.; and Weir, L.J.: “Aerodynamic Design of a 

Dual-Flow Mach 7 Hypersonic Inlet System for a 

Turbine-Based Combined-Cycle Hypersonic Propulsion 

System,” NASA/CR—2008-215214, May 2008. 

16. Auslender, A.H., Suder, K.L., Thomas, S.R.: “An 

Overview of the NASA FAP Hypersonics Project 

Airbreathing Propulsion Research, 16th 

AIAA/DLR/DGLR International Space Planes and 

Hypersonic Systems and Technologies Conference, 

AIAA–2009–7277, October 2009. 

17. Suder, K.L., Thomas, S.R., “An Overview of the NASA 

Hypersonic Project’s Combined Cycle Engine Mode 

Transition Research Experiment,” JANNAF (Joint Army, 

Navy, NASA, Air Force) 44th CS/32th APS/26th PSHS 

Joint Meeting, La Jolla, CA, December 7–11, 2009. 
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Results of an Advanced Fan Stage Operating Over a Wide Range of Speed and Bypass Ratio 

Part I: Fan Stage Design and Experimental Results 

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14. ABSTRACT

NASA’s Fundamental Aeronautics Program is investigating turbine-based combined cycle (TBCC) propulsion systems for access to space 

because it provides the potential for aircraft-like, space-launch operations that may significantly reduce launch costs and improve safety. To 

this end, National Aeronautics and Space Administration (NASA) and General Electric (GE) teamed to design a Mach 4 variable cycle 

turbofan/ramjet engine for access to space. To enable the wide operating range of a Mach 4+ variable cycle turbofan ramjet required the 

development of a unique fan stage design capable of multi-point operation to accommodate variations in bypass ratio (10×), fan speed (7×), 

inlet mass flow (3.5×), inlet pressure (8×), and inlet temperature (3×). In this paper, NASA has set out to characterize a TBCC engine fan 

stage aerodynamic performance and stability limits over a wide operating range including power-on and hypersonic-unique “windmill” 

operation. Herein, we will present the fan stage design, and the experimental test results of the fan stage operating from 15 to 100 percent 

corrected design speed. Whereas, in the companion paper, we will provide an assessment of NASA’s APNASA code’s ability to predict the 

fan stage performance and operability over a wide range of speed and bypass ratio.

15. SUBJECT TERMS

Fan stage; Compressor rotor 

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