Ashlie B. McVetta

Glenn Research Center, Cleveland, Ohio

Paul W. Giel

Vantage Partners, LLC, Brook Park, Ohio

Gerard E. Welch

Glenn Research Center, Cleveland, Ohio

Aerodynamic Investigation of Incidence Angle

Effects in a Large Scale Transonic Turbine Cascade

NASA/TM—2013-218070/REV1 AIAA–2012–3879

June 2014

This Revised Copy, numbered as NASA/TM—2013-218070/REV1, June 2014, supersedes the previous 

version, NASA/TM—2013-218070, August 2013, in its entirety. 
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Ashlie B. McVetta

Glenn Research Center, Cleveland, Ohio

Paul W. Giel

Vantage Partners, LLC, Brook Park, Ohio

Gerard E. Welch

Glenn Research Center, Cleveland, Ohio

Aerodynamic Investigation of Incidence Angle

Effects in a Large Scale Transonic Turbine Cascade

NASA/TM—2013-218070/REV1 AIAA–2012–3879

June 2014

This Revised Copy, numbered as NASA/TM—2013-218070/REV1, June 2014, supersedes the previous 

version, NASA/TM—2013-218070, August 2013, in its entirety. 

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135
Acknowledgments

This work is supported under the NASA Fundamental Aeronautics Program, Subsonic Rotary Wing Project. The authors would 

like to acknowledge the assistance of Doug Thurman (ARL) with acquiring hotwire measurements.

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at the NASA Glenn Research Center.

Level of Review: This material has been technically reviewed by technical management. 

Revised Copy

This Revised Copy, numbered as NASA/TM—2013-218070/REV1, June 2014, supersedes the previous version,

NASA/TM—2013-218070, August 2013, in its entirety. 

This paper was revised to include the GE Energy Effi cient Engine (EEE) tip section blade geometry coordinates used in this 

study. The coordinates were added to Appendix B. This additional information will allow readers to model the blade and

compare CFD results with the experiment.
NASA/TM—2013-218070/REV1 1 

Aerodynamic Investigation of Incidence Angle Effects in a 

Large Scale Transonic Turbine Cascade 

Ashlie B. McVetta 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Paul W. Giel 

Vantage Partners, LLC 

Brook Park, Ohio 44142 

Gerard E. Welch 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135

Abstract 

Aerodynamic measurements showing the effects of large incidence angle variations on an HPT 

turbine blade set are presented. Measurements were made in NASA’s Transonic Turbine Blade Cascade 

Facility which has been used in previous studies to acquire detailed aerodynamic and heat transfer 

measurements for CFD code validation. The current study supports the development of variable-speed 

power turbine (VSPT) speed-change technology for the NASA Large Civil Tilt Rotor (LCTR) vehicle. In 

order to maintain acceptable main rotor propulsive efficiency, the VSPT operates over a nearly 50 percent 

speed range from takeoff to altitude cruise. This results in 50° or more variations in VSPT blade 

incidence angles. The cascade facility has the ability to operate over a wide range of Reynolds numbers 

and Mach numbers, but had to be modified in order to accommodate the negative incidence angle 

variation required by the LCTR VSPT operation. Using existing blade geometry with previously acquired 

aerodynamic data, the tunnel was re-baselined and the new incidence angle range was exercised. Midspan 

exit total pressure and flow angle measurements were obtained at seven inlet flow angles. For each inlet 

angle, data were obtained at five flow conditions with inlet Reynolds numbers varying from 6.83105

 to 

0.85105 and two isentropic exit Mach numbers of 0.74 and 0.34. The midspan flowfield measurements 

were acquired using a three-hole pneumatic probe located in a survey plane 8.6 percent axial chord 

downstream of the blade trailing edge plane and covering three blade passages. Blade and endwall static 

pressure distributions were also acquired for each flow condition. 

Nomenclature 

Cps static pressure coefficient,     Cps  P  P2 Pt,1  P2

Cpt total pressure coefficient,     Cpt  Pt,1  Pt Pt,1  P2

Cx blade axial chord [in] 

i incidence angle [deg], des i  1 1,

M Mach number 

PR Pressure Ratio, PR  Pt,1 P2
NASA/TM—2013-218070/REV1 2 

PS pressure Surface 

P area-averaged static-pressure 

Pt area-averaged total-pressure 

Re Reynolds number, Re = ρU1Cx/μ 

s blade pitch [in] 

SS suction surface 

U total velocity 

x chordwise (axial) direction 

y tangential coordinate [in] 

Zw Zweifel coefficient, Zw = 2 1  2 cos (tan 2

Cx

s

2 tan  ) 

 Relative flow angle [deg] 

2 Angle of mass-averaged velocity components 

δ Boundary layer thickness [in] 

μ Dynamic viscosity 

ρ Density 

 Loss coefficient, ( )/( ) Pt,1  Pt,2 Pt,1  P2

Subscripts 

des Design value 

1 Cascade inlet value 

2 Cascade exit value 

i Isentropic value 

Introduction 

A conceptual Large Civil Tilt Rotor (LCTR) (Ref. 1) vehicle and mission have been developed as a part 

of NASA’s on-going work to improve the capacity of airspace systems (Refs. 2 and 3). In order to minimize 

mission fuel burn, the main rotor speeds are varied from 100 percent at takeoff to 54 percent at cruise as 

required to optimize prop efficiency at cruise. The speed variation can be accomplished through several 

means, including a variable gear-ratio (two-speed) transmission (Ref. 4), which adds complexity associated 

with gear shifting and weight. Another approach is to use a variable speed power turbine (VSPT) (Refs. 5 

and 6). The VSPT will operate at a relatively fixed corrected flow and, therefore the incidence flow angle to 

the power turbine rotor blades will vary by 50° or more over the range of VSPT shaft speeds. The current 

VSPT research effort at NASA Glenn Research Center (Ref. 7) addresses the aerodynamic challenges of a 

VSPT for NASA’s LCTR. As a part of this effort, an experimental database, required for the improvement 

and validation of turbine design methods and computational tools, is being established (Ref. 7). This will 

enable the design and optimization of incidence-tolerant blade rows for future VSPT applications. 
NASA/TM—2013-218070/REV1 3 

Little detailed experimental data are available in the open literature to support the assessment of the 

impact of extreme negative incidence on blade-row performance. Two previous studies (Refs. 8 and 9) 

examined incidence loss levels of turbine rotor blades with blunt and sharp leading edges, and each study 

also varied the pressure ratio. The authors found that loss is more sensitive to positive incidence angles, 

and that loss levels were higher for the sharp leading edge profile. The ranges of incidence studied in both 

tests (25° and 30°) were at a considerably narrower range than spanned by VSPT operation. 

In addition to the large incidence variation, the cruise altitude (28 k-ft) of the LCTR requires that the 

VSPT operate with Reynolds numbers associated with transitional flows. The impact of Reynolds number 

lapse on turbine efficiency, and the allowable speed-range of operation with acceptable efficiency, is a 

key consideration in design and analysis. Similarly, few relevant data sets exist in the open literature 

related to the aerodynamic impact of Reynolds number lapse on turbine efficiency and operability. An 

experimental data set is needed in order to have increased confidence in predictions by CFD and 

empirical models for the combined aerodynamic impact of large incidence and Reynolds number. 

The NASA Transonic Turbine Blade Cascade facility is well-suited for obtaining detailed 

aerodynamic data over a wide range of Reynolds numbers, Mach numbers, and incidence angles. An 

“incidence-tolerant” (IT) blade (Ref. 10) design is currently being tested in the facility. The IT blade 

required that the cascade be rotated to incidence angles beyond its original operational range due to 

mechanical limits; therefore, several features of the facility were redesigned to extend the negative 

incidence angle range to be compatible with VSPT requirements. 

To document and verify the aerodynamic results of the new modifications, a blade used in a previous 

aerodynamic test was chosen for this study. The test blade is a scaled rotor tip section of the GE Energy 

Efficient Engine (EEE) design (Ref. 11) that has an axial chord of 5.119 in. (130.0 mm), a blade pitch of 

5.119 in. (130.0 mm), a span of 6.000 in. (152.4 mm), and a design-intent inlet flow angle of 1 = +29.7°. 

This blade geometry was considered to be a good test case because direct comparisons could be made 

between data obtained before and after the tunnel modifications. The EEE data will also provide a good 

first data set for CFD validation over a large incidence range and transitional Reynolds numbers (Ref. 12). 

Exit total pressures, flow angles, blade loading, exit static pressures, and endwall pressures were 

acquired over a wide range of Reynolds numbers, including Reynolds numbers low enough to admit 

transitional flows as determined from previous studies in the facility (Ref. 13). The test matrix (see Table 1) 

included the design inlet Reynolds number of 683,000 and points at 1/2, 1/4, and 1/8 of this value. The 

design pressure ratio for this blade is 1.441, giving an exit isentropic Mach number of M2,i = 0.74. 

TABLE 1.—NOMINAL FLOW CONDITIONS 

[See Appendix A for actual conditions.]

Inlet Reynolds number Pressure ratio Exit isentropic 

Mach number 

683,000 (Design) 1.44 0.74 

341,500 (1/2 Design) 1.44 0.74 

170,700 (1/4 Design) 1.44 0.74 

170,700 (1/4 Design) 1.08 0.34 

85,000 (1/8 Design) 1.08 0.34 

TABLE 2.—INLET FLOW ANGLES 

Inlet angle, 

β1

Incidence angle, 

β1-βdesign

58.8° 29.1° 

48.8° 19.1° 

38.8° 9.1° 

33.8° 4.1° 

18.3° −11.4° 

–2.4° −32.1° 

–16.8° −46.5° 
NASA/TM—2013-218070/REV1 4 

The 1/8 Reynolds number case would have fallen outside of the operating envelope at the design pressure 

ratio, so it was run at a reduced pressure ratio of 1.081 (M2,i = 0.34). The impact of reducing the exit 

Mach number to establish the lowest Reynolds number was assessed, in terms of loading and loss, during 

the course of the study. The 1/4 Reynolds number cases were run at both pressure ratios for comparison. 

It is recognized that losses would scale more appropriately with exit Reynolds number, but measurements 

of the current study were based on the inlet Reynolds number for consistency with previous measurements on this same blade that focused on tip surface heat transfer and aerodynamics. The ratio of 

ReCx,2 ReCx,1 varied with inlet flow angle and PR. All values are listed in Appendix A. As listed in 

Table 2, seven incidence angles were tested ranging from +29.1° to −46.5°. The +29.1° and +9.1° 

incidence cases were repeated to re-baseline the tunnel modifications. Along with the tunnel re-baselining, 

the data from this current study will also provide a valuable and challenging experimental data set for 

CFD validation and will provide insight on the expected trends of Reynolds number and incidence angle 

variation. 

Facility Description and Modification Details 

Facility Description 

The Transonic Turbine Blade Cascade has been used for a wide range of turbine blade tests, including 

the aerodynamic tests that produced detailed flow field data for CFD code validation (Ref. 14). The 

cascade, shown in Figure 1, was described in detail by Verhoff et al. (Ref. 15) and early modifications 

that were made to the inlet to improve the inlet flow uniformity were described by Giel et al. (Ref. 16). A 

wide range of engine relevant inlet Reynolds numbers and exit Mach numbers can be achieved as seen in 

the facility operating envelope shown in Figure 2. Clean, dry, ambient temperature air is supplied to the 

cascade and the inlet pressure can be independently varied up to 14.7 psia under current facility restrictions as seen by the red dashed line. Under normal operation, a maximum inlet pressure of 23 psia 

can be achieved. The air is exhausted through a center-wide altitude exhaust system in which the pressure 

is maintained at 2 psia. The facility can run continuously at mass flows up to 58 lbm/s. The cascade test 

section nominally comprises eleven blade passages. The blades are attached to a rotating disc to allow for 

incidence angle adjustment. Upper and lower flow boards are used to direct the air from an upstream 

contraction section to the airfoils. The upper flow board is pinned to the rotating disc while the lower 

board moves independently. The boards are adjusted to be horizontal while their hinged leading edges are 

maintained in the same plane normal to the inlet flow. 

Figure 1.—Cascade Layout With EEE Blades at β1 = 38.8° (i = +9.1°). 
NASA/TM—2013-218070/REV1 5 

Figure 2.—Operating Envelope of NASA Transonic Turbine 

Cascade. 

Facility Modifications 

Under the original facility configuration, the rotating disc (Fig. 1) allowed a 45° variation in inlet 

flow angles ranging from +33.8° to +78.6° as measured from axial. This range was suitable when testing 

typical high pressure turbine blade designs but testing VSPT blade designs requires lower inlet flow 

angles for more negative incidence. The previous facility configuration, shown in Figure 3, was restricted 

to smaller negative incidence angles due to limitations on the hinge mechanism of the upper extension 

board, which can be seen in more detail in Figure 1. To solve this issue the upper board was moved to the 

minimum inlet angle limit, disconnected from the disc, and pinned to the fixed sidewall, making it 

independent of the test section. This allows the cascade to move independently of the two inlet boards. 

Since the upper board is fixed in place as the cascade is rotated in the negative incidence direction, the 

gap between the first blade and the end of the upper board increases, requiring discrete extension boards 

for each inlet angle to be tested. For this study, three extension boards were used to test the −11.4°, 

−32.1°, and −46.5° EEE incidence angles. 

Other factors which impacted the aerodynamic quality of the cascade limited the cascade’s ability to 

rotate to negative angles. Under the original facility configuration, shown in Figure 3, the exhaust duct 

extended over three-quarters of the top of the facility, leaving the upper-right corner blocked off. Figure 4 

shows the cascade section with a possible candidate IT blade design at a large negative incidence angle. 

In the configuration shown, it is evident that the airflow would be directed towards the upper-right corner 

of the cascade which, in the old configuration, is substantially blocked. As shown in Figure 4, a new 

exhaust duct was built that extends across the entire top section of the facility. The test section also had 

three large supporting blocks in the flow path around the disc, shown in Figure 3, which were needed to 

support the sidewalls under pressure or vacuum conditions. Two of the smaller blocks caused interference 

with the lower board when rotating to the larger inlet angles. The large block, located downstream of the 

blade cascade, had the potential to cause detrimental blockage when lower exit flow angle blades are 

installed. All three blocks were removed and replaced with seven 0.812 in. diameter bars located around 

the circumference of the disc. The bars are easily removable to allow the disc to rotate without interfering 

with any of the flow boards. 

The facility modifications allow for a range of inlet angles 17 78.6 . 1        The facility is shown 

in Figure 1 with the cascade at the EEE baseline inlet flow angle and the most positive (+29.1°) and 

negative incidence (−46.5°) angles that were tested. 
NASA/TM—2013-218070/REV1 6 

 

Figure 3.—Facility Before Modifications. Figure 4.—Facility After Modifications. 

Measurements 

Downstream total-pressure and exit flow angle data were obtained using a three-hole pneumatic probe 

shown in Figure 5. The probe is a 45 forward-facing pyramid probe with the measurement ports located 

on the shaft centerline. Probe measurement methods and calibration procedures similar to those described 

in Giel et al. (Ref. 14) were used. The time response on a comparable probe was measured to be to be 

0.42 sec including all tubing and pressure transducers. A 3 sec delay between the time the probe reaches 

the desired survey location and when data are recorded is imposed to ensure a 99 percent time recovery. 

The probe pitch angle coefficient was monitored to ensure that it remained within the angular calibration 

range. If needed, the probe survey was stopped and the probe was approximately nulled before resuming. 

The probe was installed in a survey plane 8.6 percent axial chords downstream of the blades that covered 

three blade passages (Station 2 in Fig. 6). The probe can traverse in a spanwise direction but for this study 

only midspan survey data were obtained. The survey consisted of 104 points over the three passages. 

Consistent with the uncertainty reported by Giel et al., (Ref. 14) the overall estimated uncertainty in flow 

angle was 1.5 and the overall estimated uncertainty in total-pressure coefficient was 1.7 percent. 

Three blades were instrumented with pressure static taps to measure blade loading. The main 

measurement blade, blade 5, was instrumented with taps at 5.5, 25, 50, 75, and 97.5 percent of span. To 

verify periodicity redundant 50, 75, and 97.5 percent, taps were placed on the suction side of blade 4 and 

blade 6 was instrumented with 50 and 97.5 percent span taps on its pressure side. Because of the midspan 

focus of the current study, only data from the 25, 50, and 75 percent span taps will be presented. The test 

section endwall is fully instrumented with 87 pressure static taps located in each blade passage and both 

upstream and downstream of the blade row. The endwall tap locations can be seen on Figure 6. Twelve 

exit static taps located approximately three and half axial chords downstream of the blades spanning over 

nearly nine pitches are used to set the exit flow condition. Inlet total pressure and temperature were 

measured with two inlet combination Kiel/total-temperature probes located on either side of the primary 

measurement passages approximately one axial chord upstream of the blades. Up to six inlet static 

pressure taps, depending on inlet flow angle, were used to determine the inlet static pressure at that same 

axial location (see Fig. 6). As seen in Figure 2, the facility has a long section upstream of the cascade that 

generates relatively thick endwall inlet boundary layers. The inlet boundary-layer thickness is estimated 

by Reynolds number scaling of detailed inlet boundary-layer measurements acquired during a previous 

study (Ref. 14). For the current study, the estimated boundary-layer thickness for each test condition can 

be found in the Table 3 of Appendix A. 
NASA/TM—2013-218070/REV1 7 

Inlet turbulence intensity measurements were made with a constant temperature single-wire hotwire 

probe that was installed in Station 0, located approximately one axial chord upstream of the blades 

(see Fig. 6). The data were acquired at a rate of 92 kHz for 7 sec. The uncertainties in the measurements 

were calculated (Ref. 17) to be less than 4 percent for the mean velocities and 5 percent for the 

fluctuations. All flow measurements of the current study were made without an inlet turbulence 

generating grid. The turbulence intensities, calculated as the normalized root-mean square of the velocity 

fluctuations, ranged from 0.25 to 0.40 percent. 

Figure 5.—Three-Hole Probe Details. 

10 9 8 7 6 5 4 3 2 1

Ps Ps Pt Ps Ps Ps , Tt Pt, Tt Pt Ps

987654321

11

y - tangential coordinate [inches]

x - axial coordinate [inches]

35 30 25 20 15 10 5 0 -5 -10 -15 -20 -25 -30

-5

0

5

10

15

20

endwalltapsv6.lay

E3 Blades and

endwall taps

exit static pressure taps

Figure 6.—E3 Endwall Static Taps and Survey Locations. 

Station 2

Station 0
NASA/TM—2013-218070/REV1 8 

Results 

Re-baseline Measurement Results 

The cascade was re-baselined after the tunnel modifications in order to document any aerodynamic 

effects caused by the new exhaust section and the support blocks. Midspan total pressures, exit flow 

angles, endwall static pressure, blade loading, and exit static pressures were previously acquired with the 

original tunnel configuration at incidence angles of +4.1°, +9.1°, +19.1°, and +29.1°. The +9.1° and 

+29.1° incidence points were repeated with the new configuration. A comparison of the exit static 

pressure data in Figures 7 and 8 show that the exit static pressure uniformity has improved with the new 

exhaust configuration. As previously mentioned, the average of the twelve exit static pressures (Fig. 6) is 

used to set the exit Mach number. This improved uniformity will result in improved periodicity and a 

more accurate setting of exit Mach number (pressure ratio). 

Although the exit pressures improved for positive incidence, nonuniform exit static pressures are 

measured as the cascade is rotated to negative incidence angles as seen in Figures 9 and 10. Illustrated in 

Figure 4, as the cascade is rotated beyond the axial position to a negative incidence, the blade-pack and 

tunnel back wall establish a converging exhaust section. The converging section (nozzle) accelerates the 

flow field as it exhausts, establishing a negative pressure gradient. This gradient is impressed on the 

blade-pack exit as a pitchwise static pressure nonuniformity (see, for example, Fig. 9). The data of 

Figures 7 to 10 were obtained at the highest Reynolds number, but all other conditions showed very 

similar results. 

5 4 6

30 20 10 0 -10 -20 -30 0.55

0.60

0.65

0.70

0.75

exit Ps(avg)

exit Ps taps

blade base pressures

endwall static taps at x = 10.000

P

Pt,in

y [inches]

(avg of 3 exit Ps taps)

(3 original Ps taps)

6 5 4

30 20 10 0 -10 -20 -30 0.55

0.60

0.65

0.70

0.75

exit Ps(avg)

exit Ps taps

blade base pressures

endwall static taps at x = 10.000

new exit Ps taps at x = 18.019

P

Pt,in

y [inches]

(avg of 12 exit Ps taps)

(3 original Ps taps)

(a) (b) 

Figure 7.—Exit Pressure at +29.1 Incidence With Original Exhaust (a) and New Exhaust (b). 

6 5 4

30 20 10 0 -10 -20 -30 0.55

0.60

0.65

0.70

0.75

exit Ps(avg)

exit Ps taps

blade base pressures

endwall static taps at x = 10.000

P

Pt,in

y [inches]

(avg of 3 exit Ps taps)

(3 original Ps taps)

6 5 4

30 20 10 0 -10 -20 -30 0.55

0.60

0.65

0.70

0.75

exit Ps(avg)

exit Ps taps

blade base pressures

endwall static taps at x = 10.000

new exit Ps taps at x = 18.019

P

Pt,in

y [inches]

(avg of 12 exit Ps taps)

(3 original Ps taps)

(a) (b) 

Figure 8.—Exit Pressure at +9.1 Incidence With Original Exhaust (a) and New Exhaust (b). 

5 4 6

30 20 10 0 -10 -20 -30 0.55

0.60

0.65

0.70

0.75

exit Ps(avg)

exit Ps taps

blade base pressures

endwall static taps at x = 10.000

new exit Ps taps at x = 18.019

P

Pt,in

y [inches]

(avg of 12 exit Ps taps)

(3 original Ps taps)

5 4 6

30 20 10 0 -10 -20 -30 0.55

0.60

0.65

0.70

0.75

exit Ps(avg)

exit Ps taps

blade base pressures

endwall static taps at x = 10.000

new exit Ps taps at x = 18.019

P

Pt,in

y [inches]

(avg of 12 exit Ps taps)

(3 original Ps taps)

Figure 9.—Exit Pressure at –11.4 Incidence. Figure 10.—Exit Pressure at –46.5 Incidence. 
NASA/TM—2013-218070/REV1 9 

PS SS

-2.0 -1.5 -1.0 y/pitch -0.5 0.0 0.5

0.0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8 1.0 0.75

1/2 0.75

1/4 0.61

1/4 0.34

1/8 0.34

Cpt dashed = old exhaust

solid = new exhaust

PS metal angle

design exit flow angle

SS metal angle

avg metal angle

-2.0 -1.5 -1.0 y/pitch -0.5 0.0 0.5 -70

-60

-50

-40

-30

 [deg]

M2,i

Re

Redes

PS SS

-2.0 -1.5 -1.0 y/pitch -0.5 0.0 0.5

0.0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8 1.0 0.75

1/2 0.75

1/4 0.75

1/4 0.34

1/8 0.34

Cpt dashed = old exhaust

solid = new exhaust

PS metal angle

design exit flow angle

SS metal angle

avg metal angle

-2.0 -1.5 -1.0 y/pitch -0.5 0.0 0.5 -70

-65

-60

-55

-50

 [deg]

M2,i

Re

Redes

(a) (b) 

Figure 11.—Exit Total Pressure and Exit Flow Angles for (a) +29.1 and (b) +9.1 Incidence Angles. 

Exit total pressure and flow angle measurements reveal that the static-pressure nonuniformity has 

little effect on the loss and flow angle periodicity of the three measurement passages of interest. The 

midspan loss coefficients and local exit flow angles are compared at i = +29.1 at five flow conditions 

in Figure 11(a). The new cascade data are shown as solid lines and the original data are shown as the 

dashed lines. Both the total pressure and flow angle data show excellent repeatability and periodicity. 

Figure 11(b) compares the data at the +9.1 angle. The measured total pressures are highly repeatable and 

the exit total pressures and flow angles are periodic. However, the new exhaust configuration did result in 

higher exit flow angles. A detailed discussion of the exit total pressures and flow angles is provided in the 

following section. 

Impact of Incidence Angle and Reynolds Number on Exit Surveys 

The effect of Reynolds number and Mach number variation on the local exit total pressure coefficient 

(top) and the flow angle (bottom) for the +29.1° incidence angle is shown in Figure 11(a). The extreme 

positive 29.1° incidence angle constitutes the most highly loaded, highest flow-turning configuration of 

the test series. At this condition, the cross-passage pressure gradient establishes a strong secondary flow 

field, which transports low-momentum boundary-layer flow to the suction-side of the blading. Note 

(Appendix A) that the endwall boundary-layers are a significant fraction of the 6.000 in. blade span. The 

aspect ratio of the test blading is sufficiently low that the low-momentum flow collects at the midspan 

symmetry plane on the suction-side of the blades. The aerodynamic blockage associated with the lowmomentum region is susceptible to rapid growth in the negative streamswise pressure gradient, and may 

ultimately separate. The combination of the strong secondary flows and the adverse pressure gradient on 

the highly loaded suction side leads to the nonuniformity (“wake”) of shown in Figure 11(a). The wake at 

this extreme positive incidence angle extends over the majority of the blade pitch. The width and depth of 

the wake increases as the Reynolds number decreases. At the two 1/4 Reynolds number conditions, the 

maximum value of Cpt in the distortion (wake) remains constant as the exit Mach number varies. The exit 

flow angle data show that there is a corresponding decrease in flow turning as the Reynolds number 

decreases, with the exception of the 1/8 Reynolds number case which has the highest flow turning. 

Though this last case appears anomalous, the flow angle data taken before and after the facility 

modifications show consistent behavior. 

At the lower positive incidence angle of +9.1°, the aerodynamic loading of the blade, though higher 

than design-intent, is reduced relative to the +29.1 setting angle. The extensive aerodynamic-blockage on 

the blade suction-side decreases as shown in Figure 11(b). The response to the Reynolds number and 

Mach number variation is similar to that at +29.1° incidence. At this angle, the wake widths remain 

essentially unchanged as the Reynolds number is held constant and the Mach number is varied. The depth 
NASA/TM—2013-218070/REV1 10 

of the wake (maximum Cpt) varies with exit Mach number, reflective of a shift in the loading diagram 

with change in Mach number (PR). At the design Mach number the flow angles are close to the average 

metal angle and flow turning decreases as the Mach number is reduced . 

At the −11.4° incidence-angle cases, shown in Figure 12, the aerodynamic loading has decreased 

further and the wake thickens only slightly on the suction side at the 1/8 Reynolds number flow condition. 

The depth of the wake shows a similar trend and remains constant at the lower Mach number as the 

Reynolds number is varied. At the 1/2 Reynolds number case, a very slight pressure-side wake thickening 

is noted. The flow angles also follow the same trend as the previous incidence angles; although, at this 

negative incidence, the flow turns to the design exit flow angle. 

Results for the highest negative incidence, i = −46.5°, are provided in Figure 13. The pitchwise totalpressure and flow angle distributions do not vary significantly with Reynolds or Mach number. A slight 

indication of possible pressure-side separation and thickening of the wake is noted. 

The effects of incidence angle for the design Reynolds number (683,000) and design exit Mach 

number (0.74) cases at the four incidence angles shown previously are summarized in Figure 14. The 

thick wake at the extreme positive incidence angle of +29.1° leads to high midspan loss. The midspan loss 

levels decrease as the inlet angle decreases. The exit flow angles increase as the incidence angles shift 

towards a negative incidence. Likewise, similar trends are evident in Figure 15 for the 1/8 Reynolds 

number at M2,i = 0.35 condition. At this lower flow condition, more effects of incidence-angle variation 

are occurring: both positive incidence cases reflect a degree of suction-side separation; and, there is 

indication of pressure-side separation at the most extreme negative incidence angle of −46.5°. 

PS SS

-2.0 -1.5 -1.0 y/pitch -0.5 0.0 0.5

0.0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

1.0 0.74

1/2 0.74

1/4 0.74

1/4 0.33

1/8 0.33

Re

Redes

Cpt M2,i

PS metal angle

design exit flow angle

SS metal angle

avg metal angle

-2.0 -1.5 -1.0 y/pitch -0.5 0.0 0.5 -70

-65

-60

-55

-50

[deg]

PS SS

-2.0 -1.5 -1.0 y/pitch -0.5 0.0 0.5

0.0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

1.0 0.74

1/2 0.74

1/4 0.74

1/4 0.33

1/8 0.34

Cpt M2,i

Re

Redes

PS metal angle

design exit flow angle

SS metal angle

avg metal angle

-2.0 -1.5 -1.0 y/pitch -0.5 0.0 0.5 -70

-65

-60

-55

-50

[deg]

Figure 12.—Effects of Reynolds Number and Pressure 

Ratio at i = −11.4°. 

Figure 13.—Effects of Reynolds Number and Pressure 

Ratio at i = −46.5°. 

PS SS

-2.0 -1.5 -1.0 y/pitch -0.5 0.0 0.5

0.0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

i = +29.1

i = +9.1

i = 11.4

i = 46.5

Cpt

PS metal angle

design exit flow angle SS metal angle

avg metal angle

-2.0 -1.5 -1.0 y/pitch -0.5 0.0 0.5 -70

-60

-50

-40

-30

 [deg]

-2.0 -1.5 -1.0 y/pitch -0.5 0.0 0.5

0.0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8 +29.1

+9.1 11.4

46.5

Cpt

PS SS

PS metal angle

design exit flow angle SS metal angle

avg metal angle

-2.0 -1.5 -1.0 y/pitch -0.5 0.0 0.5 -70

-60

-50

-40

-30

 [deg]

Figure 14.—Effects of Inlet Flow Angle at Design Flow 

Conditions (ReCx,1 and Mai,2). 

Figure 15.—Effects of Inlet Flow Angle at ReCx,1 = 84,500 

(1⁄8 design) and Mai,2 = 0.34 (75% PRdes). 

 
NASA/TM—2013-218070/REV1 11 

Impact of Incidence Angle and Reynolds Number on Blade Loading 

The blade loadings for four incidence angles, at four flow conditions are provided in Figure 16. At the 

+29.1° incidence angle, three-dimensional effects are evidenced by the spanwise static pressure taps on 

the three instrumented blades. At this angle, the weak adverse pressure gradient in the latter third of chord 

is consistent with growth/collection of aerodynamic blockage on the suction side as seen in the total 

pressure measurements of Figure 12. 

i  29.1

0.0 0.2 0.4 0.6 0.8 1.0 -1.4

-1.2

-1.0

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

i  29.1

0.0 0.2 0.4 0.6 0.8 1.0

i  29.1

0.0 0.2 0.4 0.6 0.8 1.0

i  29.1

0.0 0.2 0.4 0.6 0.8 1.0

i  9.1

0.0 0.2 0.4 0.6 0.8 1.0 -1.4

-1.2

-1.0

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

i  9.1

0.0 0.2 0.4 0.6 0.8 1.0

i  9.1

0.0 0.2 0.4 0.6 0.8 1.0

i  9.1

0.0 0.2 0.4 0.6 0.8 1.0

i  11.4

0.0 0.2 0.4 0.6 0.8 1.0 -1.4

-1.2

-1.0

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

i  11.4

0.0 0.2 0.4 0.6 0.8 1.0

i  11.4

0.0 0.2 0.4 0.6 0.8 1.0

i  11.4

0.0 0.2 0.4 0.6 0.8 1.0

i  46.5

0.0 0.2 0.4 0.6 0.8 1.0 -1.4

-1.2

-1.0

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

i  46.5

0.0 0.2 0.4 0.6 0.8 1.0

i  46.5

0.0 0.2 0.4 0.6 0.8 1.0

i  46.5

0.0 0.2 0.4 0.6 0.8 1.0

blade 6, 50% span

blade 4, 50% span

blade 4, 75% span

blade 5, 25% span

blade 5, 50% span

blade 5, 75% span

Figure 16.—Blade Loading as a Function of Incidence Angle, Mach Number, and Reynolds Number. 

x/Cx x/Cx x/Cx x/Cx 

Cps 

Cps 

Cps 

Cps 

Redes, M2,i = 0.74 Re1/4, M2,i = 0.74 Re1/4, M2,i = 0.35 Re1/8, M2,i = 0.35 
NASA/TM—2013-218070/REV1 12 

Each incidence angle shows reasonable periodicity between the three instrumented blades. The blade 

loading increases slightly due to the Mach number variation between the two 1/4 Reynolds number cases 

for each angle tested. At all angles except the +29.1° case, a comparison of the 1/4 Re cases shows that 

the location of the minimum suction surface pressure moves forward as the exit Mach number is reduced. 

The net aerodynamic loading and turning of the negative incidence cases is lower, particularly at 

i = −46.5°, relative to the positive incidence angles. The large negative incidence causes the pressure 

surface and suction surface blade loadings to be inverted near the leading edge. The magnitude of the 

inversion is larger at the lower exit Mach numbers. At a fixed negative incidence, the loading diagram 

remains fairly constant as Reynolds number is varied at a fixed Mach number (PR), but shifts noticeably 

as Mach number is changed at a fixed Reynolds number.

Impact of Incidence Angle and Reynolds Number on 

Integrated Losses and Flow Exit Angles 

The loss coefficient, , based on area-averaged total-pressures measured at the seven incidence 

angles and at each flow condition are provided in Figure 17. The integrations were calculated separately 

over two complete passages 4 and 5. As expected, losses increase at extreme negative and positive 

incidence, forming a “loss bucket”. Loss levels increase, and the loss bucket narrows, with decreasing 

Reynolds number. Incidence tolerance increases with increasing Reynolds number. 

The flow angles based on mass-weighted averaged velocity components are provided in Figure 18. 

The data reflect that the flow turning decreases with decreasing Reynolds number and with increasing 

aerodynamic loading (or increasingly positive incidence). The data asymptotically approach the design 

exit flow angle as the incidence is reduced to increasingly negative values. 

Loss was correlated with aerodynamic loading and Reynolds number. The losses are expected to be 

proportional to the square of the aerodynamic loading—here the incompressible Zweifel coefficient is 

used—and with the turbulent –1/5 power law of a turbulent flow; i.e., 2 0.2   Zw  Re . The correlation 

with data is shown in Figure 19. Data at the extreme negative incidence point, –46.5°, do not correlate 

Figure 17.—Loss Coefficient as a Function of Incidence Angle and Reynolds Number. 

Re 

Redes

 M2 Passage 
NASA/TM—2013-218070/REV1 13 

Figure 18.—Exit Flow Angle as a Function of Incidence Angle and Reynolds Number. 

Figure 19.—Re-Scaled Loss Coefficient as a Function of Incompressible Zweifel Coefficient. 

Re 

Redes

 M2 Passage 

Re 

Redes

 M2 Passage. 
NASA/TM—2013-218070/REV1 14 

with the traditional Zw coefficient, which reflects the aerodynamic loading to effect the net turning rather 

than the gross blade-row turning which is related to the combined magnitudes of the negative and positive 

constituents of the loading diagram. 

In addition to the breakdown of the adequacy of the traditional Zweifel coefficient to describe the 

combined (gross) aerodynamic loading of the negative and positive turning, the residual degree of scatter 

in Figure 19 may also be indicative of transitional flow effects in disagreement with the turbulent-flow 

scaling.

Discussion 

The experimental results described above illustrate that the EEE tip-section blading, used in the 

tunnel re-baseline effort, has a wide (60 to 70) range of incidence over which loss levels might be 

deemed acceptable. The range of useful incidence is reduced as Reynolds number decreases. Particular 

attention in VSPT design must be paid to the impact of performance and operability lapse with decreasing 

Reynolds number. This is of particular concern if the turbine design is carried out with current CFD tools 

that inadequately capture transitional flow effects. 

The loss levels and deviation angles (i.e., departure from blade exit metal angle) generally decreased 

with increasing negative incidence. Although the inlet air angles are well off-design at high negative 

incidence, the blade rows are unloaded, secondary flow fields are weak, and the flow is substantially twodimensional and follows the blading. This corroborates CFD findings at 2-D and 3-D level (Welchf 

(Refs. 6 and 18) and Suchezky (Ref. 19)) and recommends that particular attention be given to VSPT 

operating points at which aerodynamic loading is high and Reynolds numbers are low—like the LCTR 

28 k-ft cruise condition (Ref. 6)—while concurrently assuring that the lower loss levels expected for high 

negative incidence levels (e.g., –50°) at the higher Reynolds numbers takeoff condition are realized The 

VSPT will be expected to have minimal ( 15°) positive incidence-angle range at low Reynolds number 

conditions, while being able to support large ranges of negative incidence. 

No inlet turbulence grid was used in the present study. As stated, that choice was based on the intent 

to admit transitional flow at the higher Reynolds numbers in the tunnel relative to those of LCTR VSPT 

application (Ref. 7). The low freestream turbulence levels may be expected to impact loss levels at the 

negative and positive extremes of the loss bucket, and wherever flow separation and/or transition occurs. 

Further, the large inlet boundary-layer thickness (relative to blade span) and low aspect ratio of the 

cascade allows low-momentum flow to accumulate at midspan under conditions of high flow turning: the 

midspan loss levels and departure angles at the positive incidence portions of Figures 17 and 18 may be 

overly pessimistic, in that, reflecting the transport in the 3-D flow field, may be considerably higher than 

the profile losses of a 2-D section. 

Conclusions 

The results of the re-baseline measurements described in this paper demonstrate successful 

modifications of the exhaust duct and support bars in the NASA Transonic Turbine Cascade Facility to 

accommodate lower inlet flow angles. The facility is now better suited for the large incidence range 

measurements needed for the IT blade tests needed for VSPT technology development. The addition of 

twelve exit static pressure taps allows for increased accuracy in setting the exit Mach number and enables 

verification of the exit static pressure uniformity with increased accuracy. 

A key finding of the study was the strong correlation of blade-row loss levels with Zweifel 

coefficient, along with the finding that the correlation breaks down at the extreme negative incidence. The 

loss levels correlated consistently with incidence angle and Reynolds number, and the loss bucket 

narrowed (incidence-tolerance diminished) at lower Reynolds numbers. 
NASA/TM—2013-218070/REV1 15 

The data presented in this work support the idea that particular attention must be paid to the 

aerodynamic loading level and Reynolds number at the more highly loaded operating points, such as 

LCTR VSPT cruise operation, at the cost of accepting performance at higher negative incidence 

operation, such as LCTR VSPT operation at takeoff, where the blade rows are unloaded and the Reynolds 

numbers are higher. 

The detailed exit total pressures, flow angles, and blade loading data acquired over a wide range of 

Reynolds numbers, Mach numbers, and incidence angles show good repeatability, periodicity, and 

consistency with scaling laws. Those factors, along with high spatial resolution and good boundary 

characterization make this a valuable and challenging data set for CFD validation. 

NASA/TM—2013-218070/REV1 17

Appendix A—Test Matrix 

TABLE 3.—TEST MATRIX OF INLET ANGLES 

AND ACTUAL FLOW CONDITIONS 

β1 ReCx,1 ReCx,2 PR M 2,i Est. 1 [in] 

 58.8° 

(i= +29.1°)

685,000 940,700 1.4400 0.741 1.27 

341,400 469,800 1.4383 0.740 1.40 

169,800 232,000 1.2740 0.599 1.55 

171,800 230,100 1.0799 0.333 1.55 

90,300 125,700 1.0802 0.334 1.70 

 48.8° 

(i= +19.1°) 

684,700 1,122,000 1.4410 0.742 1.27 

342,900 561,400 1.4440 0.744 1.40 

170,000 274,700 1.3596 0.677 1.55 

171,800 266,100 1.0806 0.335 1.55 

86,300 136,200 1.0811 0.336 1.71 

 38.8° 

(i= +9.1°) 

685,300 1,251,000 1.4472 0.746 1.27 

343,600 628,500 1.4427 0.743 1.40 

168,200 307,300 1.4169 0.724 1.55 

173,000 296,800 1.0813 0.336 1.55 

84,500 149,000 1.0790 0.331 1.71 

 33.8° 

(i= +4.1°) 

685,100 1,297,000 1.4405 0.741 1.27 

342,300 653,300 1.4449 0.745 1.40 

169,900 323,600 1.4164 0.723 1.55 

170,500 302,300 1.0813 0.336 1.55 

87,700 158,200 1.0803 0.334 1.70 

 18.3° 

(i= −11.4.1°) 

685,100 1,570,000 1.4438 0.744 1.27 

343,300 783,300 1.4437 0.744 1.40 

170,400 387,700 1.4406 0.741 1.55 

171,300 364,400 1.0805 0.334 1.55 

84,630 183,800 1.0804 0.334 1.71 

   −2.4° 

(i= −32.1°) 

684,800 1,649,000 1.4408 0.742 1.27 

343,100 827,700 1.4443 0.744 1.40 

170,900 410,000 1.4407 0.742 1.55 

178,200 387,600 1.0812 0.336 1.54 

84,750 193,800 1.0795 0.332 1.71 

  −16.8° 

(i= −46.5°) 

685,800 1,690,000 1.4401 0.741 1.27 

343,500 841,000 1.4390 0.740 1.40 

171,700 418,600 1.4420 0.742 1.55 

172,000 385,900 1.0800 0.333 1.55 

84,760 192,000 1.0806 0.335 1.71 

NASA/TM—2013-218070/REV1 19

Appendix B—GE Energy Efficient Engine (EEE) 

Tip Section Blade Geometry Coordinates 

GE, EEE CW-22 Tip Section Blade 

Geometry 

From 2/19/2004 Coordinates 

points added, 10 smoothing iterations 

Leading Edge 

Circle: 

xc = 0.2282 yc = 0.90757 rad = 0.22598 

Trailing Edge 

Circle: 

xc = 5.03309 yc = -3.54050 rad = 0.08600 

Blade pitch = 5.1190 

X [in.] Y [in.] (blade 5) s [in.] Y [in.] (blade 4) Y [in.] (blade 6) 

5.07536 -3.61537 -7.11911 -8.73437 1.50363

5.06507 -3.62031 -7.10770 -8.73931 1.49869

5.05423 -3.62384 -7.09630 -8.74284 1.49516

5.04302 -3.62590 -7.08490 -8.74490 1.49310

5.03166 -3.62646 -7.07353 -8.74546 1.49254

5.02029 -3.62552 -7.06212 -8.74452 1.49348

5.00919 -3.62310 -7.05076 -8.74210 1.49590

4.99843 -3.61918 -7.03931 -8.73818 1.49982

4.98841 -3.61387 -7.02796 -8.73287 1.50513

4.97920 -3.60726 -7.01663 -8.72626 1.51174

4.97087 -3.59956 -7.00529 -8.71856 1.51944

4.96326 -3.59109 -6.99390 -8.71009 1.52791

4.95488 -3.58048 -6.98038 -8.69948 1.53852

4.93759 -3.55672 -6.95100 -8.67572 1.56228

4.91271 -3.52196 -6.90825 -8.64096 1.59704

4.88751 -3.48717 -6.86529 -8.60617 1.63183

4.86195 -3.45203 -6.82184 -8.57103 1.66697

4.83699 -3.41754 -6.77926 -8.53654 1.70146

4.81181 -3.38248 -6.73610 -8.50148 1.73652

4.71437 -3.24558 -6.56806 -8.36458 1.87342

4.61691 -3.10776 -6.39926 -8.22676 2.01124

4.51215 -2.95879 -6.21715 -8.07779 2.16021

4.42144 -2.82922 -6.05898 -7.94822 2.28978

4.32385 -2.68957 -5.88861 -7.80857 2.42943

4.22611 -2.54990 -5.71814 -7.66890 2.56910

4.12863 -2.41123 -5.54864 -7.53023 2.70777

4.03100 -2.27330 -5.37965 -7.39230 2.84570

3.93347 -2.13681 -5.21189 -7.25581 2.98219

3.83596 -2.00210 -5.04559 -7.12110 3.11690

3.73725 -1.86776 -4.87889 -6.98676 3.25124

3.64079 -1.73837 -4.71751 -6.85737 3.38063

3.54340 -1.60934 -4.55584 -6.72834 3.50966

3.44609 -1.48212 -4.39567 -6.60112 3.63688

3.34662 -1.35445 -4.23383 -6.47345 3.76455
NASA/TM—2013-218070/REV1 20

3.24931 -1.23256 -4.07786 -6.35156 3.88644

3.15319 -1.11526 -3.92621 -6.23426 4.00374

3.05518 -0.99848 -3.77375 -6.11748 4.12052

2.95706 -0.88427 -3.62318 -6.00327 4.23473

2.85944 -0.77357 -3.47558 -5.89257 4.34543

2.76204 -0.66663 -3.33094 -5.78563 4.45237

2.66434 -0.56346 -3.18884 -5.68246 4.55554

2.56668 -0.46475 -3.04999 -5.58375 4.65425

2.46906 -0.37068 -2.91443 -5.48968 4.74832

2.37148 -0.28122 -2.78204 -5.40022 4.83778

2.27395 -0.19628 -2.65271 -5.31528 4.92272

2.17663 -0.11597 -2.52653 -5.23497 5.00303

2.08154 -0.04185 -2.40597 -5.16085 5.07715

1.98167 0.03132 -2.28216 -5.08768 5.15032

1.88717 0.09622 -2.16752 -5.02278 5.21522

1.78625 0.16110 -2.04754 -4.95790 5.28010

1.68827 0.21999 -1.93322 -4.89901 5.33899

1.59036 0.27506 -1.82089 -4.84394 5.39406

1.49271 0.32647 -1.71053 -4.79253 5.44547

1.39587 0.37423 -1.60256 -4.74477 5.49323

1.29779 0.41937 -1.49459 -4.69963 5.53837

1.20010 0.46077 -1.38849 -4.65823 5.57977

1.10238 0.49808 -1.28389 -4.62092 5.61708

1.00483 0.53072 -1.18102 -4.58828 5.64972

0.90808 0.55830 -1.08042 -4.56070 5.67730

0.80957 0.58142 -0.97923 -4.53758 5.70042

0.71196 0.59947 -0.87997 -4.51953 5.71847

0.61435 0.61313 -0.78140 -4.50587 5.73213

0.52121 0.62310 -0.68774 -4.49590 5.74210

0.42782 0.63205 -0.59392 -4.48695 5.75105

0.33486 0.64280 -0.50034 -4.47620 5.76180

0.24179 0.65915 -0.40584 -4.45985 5.77815

0.22726 0.66267 -0.39089 -4.45633 5.78167

0.21314 0.66650 -0.37626 -4.45250 5.78550

0.19903 0.67080 -0.36151 -4.44820 5.78980

0.18535 0.67549 -0.34705 -4.44351 5.79449

0.17188 0.68066 -0.33262 -4.43834 5.79966

0.15890 0.68622 -0.31850 -4.43278 5.80522

0.14589 0.69241 -0.30409 -4.42659 5.81141

0.13334 0.69906 -0.28989 -4.41994 5.81806

0.12117 0.70620 -0.27577 -4.41280 5.82520

0.10929 0.71392 -0.26161 -4.40508 5.83292

0.09777 0.72219 -0.24743 -4.39681 5.84119

0.08663 0.73103 -0.23321 -4.38797 5.85003

0.07587 0.74050 -0.21888 -4.37850 5.85950

0.06554 0.75068 -0.20437 -4.36832 5.86968

0.05567 0.76165 -0.18961 -4.35735 5.88065

0.04636 0.77343 -0.17460 -4.34557 5.89243

0.03765 0.78613 -0.15920 -4.33287 5.90513

0.01827 0.82397 -0.11669 -4.29503 5.94297

0.00655 0.86197 -0.07692 -4.25703 5.98097

0.00078 0.90032 -0.03814 -4.21868 6.01932
NASA/TM—2013-218070/REV1 21

0.00000 0.93845 0.00000 -4.18055 6.05745

0.00062 0.95029 0.01186 -4.16871 6.06929

0.00161 0.96196 0.02357 -4.15704 6.08096

0.00296 0.97358 0.03527 -4.14542 6.09258

0.00468 0.98524 0.04706 -4.13376 6.10424

0.00674 0.99693 0.05892 -4.12207 6.11593

0.00908 1.00856 0.07079 -4.11044 6.12756

0.01167 1.02022 0.08273 -4.09878 6.13922

0.01453 1.03189 0.09475 -4.08711 6.15089

0.01766 1.04356 0.10683 -4.07544 6.16256

0.02107 1.05521 0.11897 -4.06379 6.17421

0.02475 1.06690 0.13122 -4.05210 6.18590

0.03279 1.09039 0.15605 -4.02861 6.20939

0.04995 1.13539 0.20421 -3.98361 6.25439

0.07523 1.19276 0.26690 -3.92624 6.31176

0.10041 1.24172 0.32196 -3.87728 6.36072

0.12538 1.28420 0.37124 -3.83480 6.40320

0.15006 1.32187 0.41627 -3.79713 6.44087

0.17770 1.36041 0.46370 -3.75859 6.47941

0.21486 1.40779 0.52391 -3.71121 6.52679

0.25127 1.45011 0.57974 -3.66889 6.56911

0.28739 1.48853 0.63247 -3.63047 6.60753

0.32361 1.52388 0.68308 -3.59512 6.64288

0.37229 1.56711 0.74819 -3.55189 6.68611

0.42033 1.60570 0.80980 -3.51330 6.72470

0.46921 1.64137 0.87032 -3.47763 6.76037

0.54164 1.68826 0.95660 -3.43074 6.80726

0.61437 1.72870 1.03982 -3.39030 6.84770

0.72136 1.77715 1.15726 -3.34185 6.89615

0.80965 1.80804 1.25080 -3.31096 6.92704

0.90703 1.83347 1.35145 -3.28553 6.95247

1.00486 1.85066 1.45078 -3.26834 6.96966

1.10259 1.86007 1.54896 -3.25893 6.97907

1.20045 1.86209 1.64684 -3.25691 6.98109

1.29811 1.85696 1.74464 -3.26204 6.97596

1.39527 1.84494 1.84253 -3.27406 6.96394

1.49288 1.82601 1.94196 -3.29299 6.94501

1.59093 1.80017 2.04336 -3.31883 6.91917

1.68817 1.76779 2.14585 -3.35121 6.88679

1.78572 1.72853 2.25101 -3.39047 6.84753

1.88330 1.68238 2.35895 -3.43662 6.80138

1.98104 1.62916 2.47024 -3.48984 6.74816

2.07856 1.56894 2.58486 -3.55006 6.68794

2.17623 1.50141 2.70360 -3.61759 6.62041

2.27385 1.42661 2.82658 -3.69239 6.54561

2.37150 1.34443 2.95421 -3.77457 6.46343

2.46914 1.25491 3.08667 -3.86409 6.37391

2.56678 1.15812 3.22416 -3.96088 6.27712

2.66438 1.05423 3.36670 -4.06477 6.17323

2.76203 0.94325 3.51453 -4.17575 6.06225

2.85958 0.82550 3.66744 -4.29350 5.94450

2.95717 0.70097 3.82565 -4.41803 5.81997
NASA/TM—2013-218070/REV1 22

3.05488 0.56968 3.98931 -4.54932 5.68868

3.15246 0.43209 4.15799 -4.68691 5.55109

3.25002 0.28821 4.33183 -4.83079 5.40721

3.34773 0.13795 4.51107 -4.98105 5.25695

3.44530 -0.01815 4.69514 -5.13715 5.10085

3.54290 -0.18028 4.88438 -5.29928 4.93872

3.64053 -0.34844 5.07883 -5.46744 4.77056

3.73816 -0.52253 5.27843 -5.64153 4.59647

3.83582 -0.70254 5.48322 -5.82154 4.41646

3.93353 -0.88858 5.69336 -6.00758 4.23042

4.03068 -1.07950 5.90757 -6.19850 4.03950

4.12869 -1.27811 6.12905 -6.39711 3.84089

4.22614 -1.48131 6.35442 -6.60031 3.63769

4.32369 -1.69000 6.58477 -6.80900 3.42900

4.42143 -1.90392 6.81997 -7.02292 3.21508

4.51893 -2.12168 7.05857 -7.24068 2.99732

4.61672 -2.34408 7.30151 -7.46308 2.77492

4.71446 -2.56988 7.54756 -7.68888 2.54912

4.81186 -2.79773 7.79535 -7.91673 2.32127

4.86193 -2.91561 7.92343 -8.03461 2.20339

4.91216 -3.03423 8.05225 -8.15323 2.08477

4.96207 -3.15234 8.18046 -8.27134 1.96666

5.01197 -3.27061 8.30883 -8.38961 1.84839

5.06204 -3.38896 8.43733 -8.50796 1.73004

5.09250 -3.45994 8.51458 -8.57894 1.65906

5.10125 -3.47983 8.53631 -8.59883 1.63917

5.10768 -3.49485 8.55264 -8.61385 1.62415

5.11235 -3.50688 8.56555 -8.62588 1.61212

5.11577 -3.51774 8.57694 -8.63674 1.60126

5.11809 -3.52881 8.58824 -8.64781 1.59019

5.11900 -3.54010 8.59958 -8.65910 1.57890

5.11836 -3.55150 8.61099 -8.67050 1.56750

5.11613 -3.56281 8.62252 -8.68181 1.55619

5.11252 -3.57341 8.63372 -8.69241 1.54559

5.10746 -3.58365 8.64514 -8.70265 1.53535

5.10109 -3.59311 8.65655 -8.71211 1.52589

5.09357 -3.60162 8.66790 -8.72062 1.51738

5.08495 -3.60908 8.67930 -8.72808 1.50992

5.07536 -3.61537 8.69077 -8.73437 1.50363
NASA/TM—2013-218070/REV1 23

References 

1. Johnson, W., Yamauchi, G.K., and Watts, M.E., “NASA Heavy Lift Rotorcraft Systems

Investigation,” NASA/TP—2005-213467, Sep., 2005.

2. Acree, C.W., Hyeonsoo, Y., and Sinsay, J.D., “Performance Optimization of the NASA Large Civil

Tiltrotor,” Proc. of the Int. Powered Lift Conf., Jul, 2008.

3. Wilkerson, J.B. and Smith, R.L., “Aircraft System Analysis of Technology Benefits to Civil

Transport Rotorcraft,” NASA/CR—2009-214594, Jun., 2009.

4. Stevens, M.A., Handschuh, R.F., Lewicki, D.G., “Variable/Multispeed Rotorcraft Drive System

Concepts,” NASA/TM—2009-215456; also ARL-TR-4728, Mar., 2009.

5. D′Angelo, M., “Wide Speed Range Turboshaft Study,” NASA/CR—1995-198380, Aug., 1995.

6. Welch, G.E., “Assessment of Aerodynamic Challenges of a Variable-Speed Power Turbine for a

Large Civil Tilt-Rotor Application,” Proc. of the AHS Int. 66th Annual Forum, Arizona, May 2010;

also NASA/TM—2010-216758.

7. Welch, G.E., McVetta, A.B., Stevens, M.A., Howard, S.A., Giel, P.W., Ameri, A.A., Waiming, T.,

Skoch, G.J., Thurman, D.R., “Variable-Speed Power-Turbine Research at Glenn Research Center,”

Proc. of the AHS Int.68th Annual Forum, Texas, May 2012; also NASA/TM—2012-217605.

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Cascade,” NASA TM-X-3047, 1974.

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2011-217016, 2012.

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No. 1, pp. 110-121, Jan. 2004.

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W.P., La, H., “Three Dimensional Flow Field Measurements in a Transonic Turbine Cascade,”

ASME 96-GT-113, presented at the 1996 ASME/IGTI Gas Turbine Conference, Birmingham,

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217424, Feb. 2012.
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4. TITLE AND SUBTITLE

Aerodynamic Investigation of Incidence Angle Effects in a Large Scale Transonic

Turbine Cascade

5a. CONTRACT NUMBER

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6. AUTHOR(S)

McVetta, Ashlie, B.; Giel, Paul, W.; Welch, Gerard, E.

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5f. WORK UNIT NUMBER

WBS 380046.02.03.02.01.01

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

John H. Glenn Research Center at Lewis Field

Cleveland, Ohio 44135-3191

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E-18746-1

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13. SUPPLEMENTARY NOTES

This Revised Copy, numbered as NASA/TM-2013-218070-REV1, June 2014, supersedes the previous version, NASA/TM—2013-218070,

August 2013, in its entirety.

14. ABSTRACT

Aerodynamic measurements showing the effects of large incidence angle variations on an HPT turbine blade set are presented.

Measurements were made in NASA’s Transonic Turbine Blade Cascade Facility which has been used in previous studies to acquire detailed

aerodynamic and heat transfer measurements for CFD code validation. The current study supports the development of variable-speed power

turbine (VSPT) speed-change technology for the NASA Large Civil Tilt Rotor (LCTR) vehicle. In order to maintain acceptable main rotor

propulsive efficiency, the VSPT operates over a nearly 50 percent speed range from takeoff to altitude cruise. This results in 50° or more

variations in VSPT blade incidence angles. The cascade facility has the ability to operate over a wide range of Reynolds numbers and Mach

numbers, but had to be modified in order to accommodate the negative incidence angle variation required by the LCTR VSPT operation.

Using existing blade geometry with previously acquired aerodynamic data, the tunnel was re-baselined and the new incidence angle range

was exercised. Midspan exit total pressure and flow angle measurements were obtained at seven inlet flow angles. For each inlet angle, data

were obtained at five flow conditions with inlet Reynolds numbers varying from 6.83×105

 to 0.85×105 and two isentropic exit Mach

numbers of 0.74 and 0.34. The midspan flowfield measurements were acquired using a three-hole pneumatic probe located in a survey plane

8.6 percent axial chord downstream of the blade trailing edge plane and covering three blade passages. Blade and endwall static pressure

distributions were also acquired for each flow condition.

15. SUBJECT TERMS

Turbomachinery; Turbines; Axial flow turbines; Aerodynamics

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