Ashlie B. Flegel

Glenn Research Center, Cleveland, Ohio

Paul W. Giel

Vantage Partners, LLC, Brook Park, Ohio

Gerard E. Welch

Glenn Research Center, Cleveland, Ohio

Aerodynamic Effects of High Turbulence Intensity

on a Variable-Speed Power-Turbine Blade With

Large Incidence and Reynolds Number Variations

NASA/TM—2014-218137

September 2014
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Ashlie B. Flegel

Glenn Research Center, Cleveland, Ohio

Paul W. Giel

Vantage Partners, LLC, Brook Park, Ohio

Gerard E. Welch

Glenn Research Center, Cleveland, Ohio

Aerodynamic Effects of High Turbulence Intensity

on a Variable-Speed Power-Turbine Blade With

Large Incidence and Reynolds Number Variations

NASA/TM—2014-218137

September 2014

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

50th Joint Propulsion Conference

cosponsored by AIAA, ASME, SAE, and ASEE

Cleveland, Ohio, July 28–30, 2014
Acknowledgments

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

to acknowledge Doug Thurman (Army Research Laboratory) for his assistance with hotwire measurements. The authors also 

thank Adam Ford, Matt Bloxham, Steve Gegg (deceased), and Ed Turner of Rolls-Royce North American Technologies for their 

contributions during the blade design efforts under the NASA RTAPS contract. It was the authors’ great honor to have worked 

with Dr. Steven G. Gegg.

Available from

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7115 Standard Drive

Hanover, MD 21076–1320

National Technical Information Service

5301 Shawnee Road

Alexandria, VA 22312

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

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—2014-218137 1 

Aerodynamic Effects of High Turbulence Intensity on a 

Variable-Speed Power-Turbine Blade With Large 

Incidence and Reynolds Number Variations 

Ashlie B. Flegel 

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 

The effects of high inlet turbulence intensity on the aerodynamic performance of a variable speed 

power turbine blade are examined over large incidence and Reynolds number ranges. These results are 

compared to previous measurements made in a low turbulence environment. Both high and low 

turbulence studies were conducted in the NASA Glenn Research Center Transonic Turbine Blade 

Cascade Facility. The purpose of the low inlet turbulence study was to examine the transitional flow 

effects that are anticipated at cruise Reynolds numbers. The current study extends this to LPT-relevant 

turbulence levels while perhaps sacrificing transitional flow effects. Assessing the effects of turbulence at 

these large incidence and Reynolds number variations complements the existing database. Downstream 

total pressure and exit angle data were acquired for 10 incidence angles ranging from +15.8° to −51.0°. 

For each incidence angle, data were obtained at five flow conditions with the exit Reynolds number 

ranging from 2.12×105

 to 2.12×106 and at a design exit Mach number of 0.72. In order to achieve the 

lowest Reynolds number, the exit Mach number was reduced to 0.35 due to facility constraints. The inlet 

turbulence intensity, Tu, was measured using a single-wire hotwire located 0.415 axial-chord upstream of 

the blade row. The inlet turbulence levels ranged from 8 to 15 percent for the current study. Tu

measurements were also made farther upstream so that turbulence decay rates could be calculated as 

needed for computational inlet boundary conditions. Downstream flow field measurements were obtained 

using a pneumatic five-hole pitch/yaw probe located in a survey plane 7 percent axial chord aft of the 

blade trailing edge and covering three blade passages. Blade and endwall static pressures were acquired 

for each flow condition as well. The blade loading data show that the suction surface separation that was 

evident at many of the low Tu conditions has been eliminated. At the extreme positive and negative 

incidence angles, the data show substantial differences in the exit flow field. These differences are 

attributable to both the higher inlet Tu directly and to the thinner inlet endwall boundary layer that the 

turbulence grid imposes. 

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.] 
NASA/TM—2014-218137 2 

H blade span [in.] 

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

iopt incidence angle with minimum loss [deg], iopt = –26.0° 

M Mach number 

n Exponent on Reynolds number scaling (n = 0.5 laminar, 0.1 turbulent) 

PR Pressure Ratio, PR  Pt,1 P2

PS Pressure Surface 

P area-averaged static-pressure 

Pt area-averaged total-pressure 

Re exit isentropic Reynolds number, Re = ρ2,iU2,iCx / μ2

Reb baseline Reynolds number, Reb = 5.30×105 

S blade pitch [in.] 

SS suction surface 

Tu turbulence intensity, Tu u U 2 

U total velocity [ft/s] 

U mean velocity [ft/s], U = (Ux, Uy, Uz) 

u fluctuating velocity [ft/s], u = (ux, uy, uz) 

x chordwise (axial) coordinate [in.] 

y pitchwise (tangential) coordinate [in.] 

z spanwise coordinate [in.] 

Zw Zweifel coefficient,   2 1 2

2 cos tan tan

2      Cx

S Zw

 relative flow angle, pitch angle [deg],  = tan–1(Uy / Ux) 

2 angle of mass-averaged exit velocity components 

Δβ2 departure angle from trailing edge mean camber line 

γ yaw angle [deg], γ = tan–1(Uz / Ux) 

δ99 boundary layer thickness [in.] 

μ dynamic viscosity 

ρ density 

ω  loss coefficient,       Pt,1  Pt Pt,1  P2

ωc corrected loss coefficient, ωc = ω (Reb/Re) –n 

Subscripts: 

1 cascade inlet value measured at Station 0 

2 cascade exit value 

i isentropic value 

s streamwise component 

t total condition 
NASA/TM—2014-218137 3 

Introduction 

Variable Speed Power Turbine (VSPT) technology has been an area of focus at NASA Glenn 

Research Center under the Fundamental Aeronautics Program Rotary Wing Project (Ref. 1). The VSPT 

will have the ability to operate over a large (e.g., 50 percent) shaft-speed range, which enables required 

main-rotor speed variation of the notional Large Civil Tiltrotor (LCTR) concept. The LCTR could help 

relieve airport congestion and increase airspace through-put capacity by utilizing its VTOL and Mach 0.5 

cruise capability (Refs. 2 and 3). In order to optimize the propulsive efficiency, the main rotors of the 

LCTR must vary from 100 percent speed at takeoff to 54 percent speed at cruise. The key aerodynamic 

challenges of the VSPT, due to this speed change, include achieving high turbine efficiency at high work 

factors, managing the loss levels over a large (40° to 60°) incidence variation, and operating at low cruise 

Reynolds numbers at which transitional flow will become a factor. Little data exist in the open literature 

that examines the effects of a large incidence or Reynolds number variation. 

A recent study was conducted on a VSPT blade geometry in NASA’s Transonic Turbine Blade 

Cascade Facility over a 66.8° incidence range covering an order of magnitude variation of Reynolds 

numbers ranging from 2.12×105

 to 2.12×106 (Ref. 4). That test was conducted at low inlet turbulence 

(Tu = 0.24 to 0.4 percent) in order to admit transitional flow on the blade surface since facility limitations 

did not allow the Reynolds number to be reduced directly to LCTR applicable values. The results of that 

study showed a significant effect of strong secondary flows associated with high aerodynamic loading 

levels at the large positive incidence angles. The secondary flows weakened with decreasing aerodynamic 

loading at negative incidence conditions and the flow became substantially more two-dimensional. At this 

low inlet Tu, the measured blade loadings reflected transitional flow on the suction surface. Inverted 

(negative) portions of the pressure loading distributions reflected negative flow turning at the higher 

negative incidence angles. Separation was also noted on the pressure surface at extreme negative 

incidence angles (Ref. 4). 

The focus of this current study is to document the aerodynamic effects of a large range of incidence 

and Reynolds number variation at engine realistic turbulence levels. An upstream blowing turbulence 

grid was installed. In addition to producing turbulence, the grid configuration reset and significantly 

(50 percent) reduced the endwall boundary layer thickness entering the blade cascade. The reduction in 

inlet boundary-layer thickness had an influence on the aerodynamic blockage levels and total-pressure 

distribution in the cascade. 

The comprehensive database and associated VSPT blade-section geometry is to be made publicly 

available. The wide incidence range and turbulence intensity range of the test matrix make the dataset of 

much value for LPT-relevant computational model assessment (e.g., see the recent study by Ameri et al. 

(Refs. 5 and 6)). 

Test Facility, Turbulence Grid, and Blade Description 

Facility 

This study of a two-dimensional VSPT blade section was conducted in the NASA Glenn Research 

Center (GRC) Transonic Turbine Blade Cascade Facility shown in Figure 1. This facility has been used in 

several turbine aerodynamic and heat transfer studies, (Refs. 7 to 10) including the assessment of the 

VSPT blade at low inlet Tu. The operating envelope shown in Figure 2 demonstrates the large 

independent range of engine relevant Reynolds and Mach numbers within facility capability. Inlet air is 

supplied by GRC’s 40 psig Combustion Air system. This clean, dry, ambient temperature air enters the 

facility at a maximum inlet pressure of 14.7 psia under current safety restrictions. This restriction is 

represented as a red dashed line in Figure 2. The air passes through a series of flow conditioning and 

contraction sections and is directed into the cascade test section. The air is then discharged into an exhaust 

header at 2 psia. Valves between the test section and the exhaust header are independently adjusted to 

give the desired Reynolds and Mach numbers. 
NASA/TM—2014-218137 4 

Figure 1.—Cascade layout with VSPT geometry at i = 5.8°. 

Figure 2.—CW-22 operating envelope. 

The cascade test section, illustrated in Figure 1, is made up of nominally 10 blade passages. The 

blades are mounted on a rotating disk that can set the inlet flow angles (from axial) in the range of −17° 

β1  +78.8°. Measurements were acquired at the 10 incidence angles listed in Table 1. Additional detailed 

measurements were obtained at two angles representing takeoff (i = +5.8°) and cruise (i = −36.7°) flight 

mission points. Unique upper flow board extensions with respective blade suction-side profiling were 

installed, replacing the first blade, for five incidence angles in the range of −16.1°  i  −51.0°. These 

extensions ensured that the flow was properly directed into the first blade passage, the upper and lower 

flow boards were horizontal, and their hinged leading edges were maintained in the same plane normal to 

the inlet flow. 

Midspan total-pressure and exit angle data were acquired for each incidence angle tested at five 

nominal flow conditions. These conditions are represented as the pink triangles in Figure 2 and are listed 

in Table 2. The design pressure ratio was 1.412, which corresponds to an isentropic exit Mach number of 

0.72. A baseline flow condition was achieved by finding the lowest exit Reynolds number at which the 

tunnel could consistently maintain the design exit Mach number. This baseline Reynolds number, Reb,

was found to be 0.53×106

. The Reynolds number was increased to 1.06×106 and 2.12×106 to examine the 

effects of higher Reynolds numbers. To reduce Reynolds number to 2.12×105

 so as to enable an order-ofmagnitude variation of Reynolds number within facility capability, the exit Mach number had to be 

reduced to M2,i = 0.35. At this exit Mach number, data were acquired at Reb and at 2.12×105

 which 

allowed the Mach number effects to be isolated at Reb. 























Ref. [4] and current study

M2,i,des=0.72

M2,i=0.35

Ref. [8]

Ref. [9]

Ref. [10]

Isentropic exit unit Reynolds number,

0.0

0.5

1.0

1.5

2.0

0.1 0.2 0.3 0.5 0.7 1.0 2 3 5 7 10

Re2,i  106 [1/ft]

Isentropic

exit Mach

number,

M2,i

minimum

exhaust

pressure 13.8 kPa

(2.0 psia)

max mass flow

 26 kg/s

(58 lbm/s)

maximum

inlet pressure

= 159 kPa

(23.0 psia)

7.0

6.0

5.0

4.0

3.0

2.0

1.5

1.3

1.2

1.1

1.0

Pressure

Ratio,

Pt,1

P2

Pt = 14.7 psia
NASA/TM—2014-218137 5 

TABLE 1.—INLET FLOW ANGLES AND 

CORRESPONDING ZWEIFEL COEFFICIENT 

Inlet angle, 

β1 

Incidence angle, 

i

Zw 

50.0° 

45.0° 

40.0° (cruise) 

34.2° 

28.0° 

18.1° 

8.2° 

−2.5° (takeoff) 

−11.8° (mission max-i) 

−16.8° 

15.8° 

10.8° 

5.8° 

0.0° 

−6.2° 

−16.1° 

−26.0° 

−36.7° 

−46.0° 

−51.0° 

1.22 

1.13 

1.06 

0.99 

0.92 

0.82 

0.74 

0.65 

0.58 

0.53 

TABLE 2.—NOMINAL FLOW CONDITIONS 

Flow parameters Low Tu1 High Tu1 

Exit ReCx Pressure 

ratio 

Exit Mis  99,1 a

[in.] 

299,1/Ha  99,1 a

[in.] 

299,1/Ha

2.12×106 (4.0Reb) 1.412 0.72 1.16 to 1.23 0.39 to 0.41 0.58 to 0.62 0.19 to 0.21 

1.06×106 (2.0Reb) 1.412 0.72 1.28 to 1.36 0.43 to 0.45 0.64 to 0.69 0.21 to 0.23 

5.30×105 (1.0Reb) 1.412 0.72 1.42 to 1.50 0.47 to 0.50 0.71 to 0.76 0.24 to 0.25 

5.30×105 (1.0Reb) 1.087 0.35 1.40 to 1.49 0.47 to 0.50 0.71 to 0.75 0.24 to 0.25 

2.12×105 (0.4Reb) 1.087 0.35 1.60 to 1.69 0.53 to 0.56 0.81 to 0.86 0.27 to 0.29 a

Reynolds-scaling estimated range of endwall boundary-layer thickness at cascade inlet over ten incidence angle 

settings.

Inlet boundary-layer measurements were made at –0.415 Cx for i = +5.8° and i = –36.7°; and over the 

current range of exit Reynolds numbers. The inlet boundary layer measurements indicated that 99,1

correlated to the inlet Reynolds number only. A boundary-layer thickness correlation was developed of 

the form 99,1  ReCx,1

–1/7 as appropriate for turbulent flow. The inlet boundary layer thickness ranges 

listed in Table 2 were calculated from the correlation for all 10 incidence angles over the corresponding 

inlet Reynolds number ranges. It is also important to note that the boundary layer thicknesses decreased 

by a factor of two for these high Tu tests. Even though the boundary layer is reduced upon grid 

installation, the endwall boundary layers still cover 20 to 30 percent of the blade span. 

Turbulence Grid Description 

The focus of this study was to examine the effects of turbulence intensity over a large range of 

Reynolds number and incidence. To generate turbulence, an upstream blowing turbulence grid was 

installed roughly five-axial chords upstream of the center of the blade row leading edge. The grid, seen in 

Figure 1, was made of 1 in. square tubes. A vertical tube spanned the upper and lower inlet board at 

midspan. Depending on the incidence angle and the distance between the upper and lower inlet boards, 

the grid consisted of five or six horizontal 1 in. square tubes located on 6 in. centers. Air entered the grid 

through both ends of the horizontal tubes and exited upstream through 1/8 in. diameter holes. The holes 

were spaced uniformly on half inch centers and centered on their respective tube. It has been found that 

this type of grid with upstream blowing produced a more uniform mean and fluctuating flowfield 

compared to an unblown grid (Ref. 11). The optimum grid blowing ratio was established by surveying the 

inlet flow at Station 0 (Fig. 3) while varying blowing flow rate until the best mean flow uniformity was 

achieved. 

A constant temperature single-wire hotwire anemometer probe was used to acquire inlet turbulence 

intensity data. The hotwire was installed in Station 0, located approximately 0.415 Cx upstream of the 

blades as shown in Figure 3. The turbulence intensity was documented to be between 8 to 15 percent, 

depending on inlet flow angle and the resultant distance between the upstream grid and cascade inlet. 

Details of the hotwire measurements can be found in Thurman et al. (Ref. 12). 
NASA/TM—2014-218137 6 

Figure 3.—VSPT blade and measurement plane details. 

Blade Description 

The blade geometry is a scaled-up two-dimensional midspan section of a second-stage rotor of the 

4-stage VSPT design (Ref. 13). Details of the blade and the design/optimization process used to establish 

the two-dimensional profile are documented elsewhere (Refs. 4 and 13). The blade geometry is shown in 

Figure 3 and details of the scaled (test) blade are described in Table 3. 

A computed design-point blade loading diagram is compared to the experimental data from the 

previous low Tu test and the current high Tu test in Figure 4. The experimental data is taken at midspan 

for two Reynolds number conditions. The design calculation (Ref. 13) was carried out on a twodimensional cone assuming fully turbulent flow. It is noted that the high Tu data correspond very well 

with the design intent throughout the front portion of the blade; however the disparity between computed 

and measured pressure coefficients on in the aft portion of the blade, on pressure side and suction side, is 

5

6

4

3

passage

4

passage

5

passage

3

Sta. 0

Sta. 2

2.0

0.0

55.5 metal

1.0

y 

pitch

1=34.2 zero i

1=11.8 max i

1=2.5 t-o

1=40.0 cruise

y

x

x/C 1.070 x x/C = x =

0.415

i=5.8

i=46.0

i=36.7

z

x

 1

Cx

H

 1
NASA/TM—2014-218137 7 

TABLE 3.—BLADE DESCRIPTION 

Parameter Value 

Axial chord, Cx

True chord 

Pitch, S 

Span, H

Throat dimension 

Leading edge diameter 

Trailing edge diameter 

Stagger angle 

Inlet metal angle 

Uncovered turning 

Exit metal angle 

180.57 mm (7.109 in.) 

194.44 mm (7.655 in.) 

130.00 mm (5.119 in.) 

152.40 mm (6.000 in.) 

72.85 mm (2.868 in.) 

15.16 mm (0.597 in.) 

3.30 mm (0.130 in.) 

20.35° 

34.2° 

19.47° 

−55.54° 

Figure 4.—Design intent and experimental data at i = +5.8°. 

evident. This may be attributed, in part, to differences in the reference exit-pressure of the pressure 

coefficients in computation and experiment. The disparity might also be associated with streamtube 

contraction in experiment due to the endwall blockage that is not present in the CFD. The low Tu data 

indicate a region of transition at x/Cx = 0.85 which is influenced by the strong three-dimensionally and 

secondary flow fields in the low aspect ratio cascade (Fig. 3). 

Measurement Description 

Total-pressure and exit flow angle data were obtained using a five-hole pitch-yaw probe installed in 

the Station 2 survey plane shown in Figure 3. The survey plane is located approximately 7.0 percent 

axial-chord downstream of the blade trailing edge and covers three blade passages. The five-hole probe is 

a 45° forward-facing pyramid probe with the measurement port located on the shaft centerline. Midspan 

surveys were obtained for all 10 incidence angles consisting of 123 pitchwise points spaced non-uniformly 

over three blade passages. For incidence angles +5.8° (cruise) and −36.7° (takeoff), additional surveys were 

acquired at 12 spanwise immersions uniformly spaced from 0.042 ≤ z/H ≤ 0.50. For these surveys, each 

spanwise point consisted of 62 pitchwise points spaced uniformly across three blade passages. 

 

x/Cx

0.0 0.2 0.4 0.6 0.8 1.0

-1.2

-1.0

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

low Tu, 1 Reb

low Tu, 4 Reb

high Tu, 1 Reb

high Tu, 4 Reb

design intent

Cps
NASA/TM—2014-218137 8 

Detailed probe descriptions and calibration methods were described in Giel et al. (Ref. 8). The time 

response of the five-hole probe was measured to be 0.42 s. A 3 sec delay was imposed between the time 

the probe reached the desired survey location and the data recording initiation to ensure a 

95 percent time recovery. During each survey, the probe was monitored to ensure it remained well within 

the angular calibration range of 40°. The overall estimated absolute uncertainty in flow angle was 1.5° 

and the overall estimated local uncertainty in total-pressure coefficient was 0.01 based on the analysis 

reported in Giel et al. (Ref. 8). 

The primary measurement blades 4, 5, and 6 were instrumented with static pressure taps at four 

spanwise locations. Blade 5 was fully instrumented on the pressure and suction surfaces with 44 taps 

located at 10, 15, 30, and 50 span. Blade 4 was instrumented with 20 redundant taps on the suction side 

and blade 6 had 16 redundant taps on the pressure side. The endwall was also fully instrumented with 

static taps located upstream and downstream of the blade row and within the blade passages. 

The exit Mach number condition was set by the average of 12 exit static-pressure taps located 

approximately three axial-chords downstream of the blades and spanning nine blade pitches. The inlet 

static pressure was measured by five to six inlet static pressure taps located 2.95 in. upstream of the blade 

row at Station 0. Inlet total pressure and temperature were measured with two combination Kiel/totaltemperature probes located upstream at midspan in Station 0, just outside the passages measured by the 

downstream survey probe. 

Results 

Impact of the Turbulence Grid and Incidence on the Three-Dimensional Flowfield 

Survey Data—Detailed flowfield surveys were acquired for the cruise (i = +5.8°, β2 = 40.0°) and 

takeoff (i = –36.7°, β2 = –2.5°) incidence angles at the baseline Reynolds number (Reb) and nominal 

design exit Mach number (M2 = 0.72). The current high Tu data were acquired with a five-hole probe over 

12 pitchwise surveys at spanwise immersions between 0.042  z/H  0.50. For the previous low inlet Tu

data, 14 additional spanwise points were acquired with a three-hole probe in the near-wall region between 

0.0  z/H  0.043. Total pressure coefficient contours for the positive incidence, i = +5.8°, are shown in 

Figure 5. At low inlet Tu (Fig. 5(a)), the blade is highly loaded producing strong secondary flows that 

transport the thick (e.g., 50 percent of the half-span) endwall flow to and along the suction side of the 

blade. A region of elevated Cpt in the wake at midspan reflects separation, transition, and reattachment at 

low Tu (also evident in Fig. 4). At the same highly loaded condition, but with the blown turbulence grid 

installed (high Tu), the impact of reduced inlet boundary layer thickness is evident in Figure 5(b) in that 

far less low total pressure flow has accumulated on the blade suction-side; furthermore, the midspan wake 

loss levels are lower, reflecting that the flow in the high Tu environment remains attached (no suctionside separation, see Fig. 4). 

At the high negative incidence condition (Fig. 6), the blade loading is reduced and the flow becomes 

largely two-dimensional. As the strength of the secondary flow is reduced at the low load condition, 

transport of the boundary layer flow up the suction-side is reduced. The thick boundary layer of the low 

Tu (Fig. 6(a)) remains near the endwall. 

All of the total-pressure contours for both inlet angles show good passage-to-passage periodicity. 

Note that the periodicity improved for the high Tu case. 

Figure 7 shows secondary flow vectors overlaid on the total-pressure contours (Fig. 7(a)), pitch angle 

contours (Fig. 7(b)), and yaw angle contours (Fig. 7(c)) at high inlet Tu for the positive incidence 

(i = +5.8°) condition over a single passage. Near the endwall (z/H = 0.04) there is evidence of strong 

overturning at the hub. The reduced inlet boundary layer of the high Tu configuration causes the 

secondary flow vectors to diminish as compared to the low Tu data shown in Figure 8(a), with less 

evidence of the horseshoe and passage vortices. Since the flow is more two-dimensional for i = –36.7°, 

there is little variation as shown in Figure 9. 
NASA/TM—2014-218137 9 

(a) Low Tu

(b) High Tu

Figure 5.—Total pressure coefficient contours at Reb and M2,i = 0.72 at i = +5.8°. 

(a) Low Tu

(b) High Tu

Figure 6.—Total pressure coefficient contours at Reb and M2,i = 0.72 at i = –36.7°. 

 

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

0.0

0.1

0.2

0.3

0.4

z 0.5 H

PS SS PS SS PS SS

Passage 4 Passage 5

-2.0 -1.5 -1.0 -0.5 0.0 0.5

0.0

0.1

0.2

0.3

0.4

0.5

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

0.0

0.1

0.2

0.3

0.4

0.5

0.00 0.10 0.20 0.30 0.40 0.50 0.60 0.70 0.80

Cpt :

z

H

PS SS PS SS PS SS

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

0.0

0.1

0.2

0.3

0.4

z 0.5 H

PS SS PS SS PS SS

-2.0 -1.5 -1.0 -0.5 0.0 0.5

0.0

0.1

0.2

0.3

0.4

0.5

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

0.0

0.1

0.2

0.3

0.4

0.5

0.00 0.10 0.20 0.30 0.40 0.50 0.60 0.70 0.80

Cpt :

z

H

PS SS PS SS PS SS
NASA/TM—2014-218137 10 

Figure 7.—Detailed view of high Tu flow at i = +5.8°, Reb, M2,i = 0.72. 

(a) (b) 

Figure 8.—Low Tu secondary flow contours at (a) i = +5.8° and (b) at i = –36.7°; Reb, M2,i = 0.72. 

Figure 9.—Detailed view contours of high Tu flow at i = –36.7°, Reb, M2,i = 0.72. 

y/S -1.0 -0.5 0.0

0.0

0.1

0.2

0.3

0.4

0.5

z

H

-1.0 -0.5 0.0

0.0

0.1

0.2

0.3

0.4

0.5

t,1 1 t,1 2

-1.0 -0.5 0.0 0.0

0.1

0.2

0.3

0.4

0.5

-58 -54

-50

-50

-46

-42

-50

-54

-54 -54

-58

-62

-50

-1.0 -0.5 0.0 0.0

0.1

0.2

0.3

0.4

0.5 alpha: -78 -74 -70 -66 -62 -58 -54 -50 -46 -42 -38 -34 -30 -26 -22

 [deg]

-1.0 -0.5 0.0 0.0

0.1

0.2

0.3

0.4

0.5

0

0

0

-2

2

4

6 4 2

0 0

-1.0 -0.5 0.0 y/S 0.0

0.1

0.2

0.3

0.4

0.5 beta: -6 -4 -2 0 2 4 6 8 10 12 14 16 18 20 22 24 26

 [deg]

z

H

(a) Total Pressure Coefficient and Secondary Flow Vectors 

(b) Pitch Angle, β [deg]

(c) Yaw Angle, γ [deg]

y/S -1.0 -0.5 0.0

0.0

0.1

0.2

0.3

0.4

0.5

z

H

-1.0 -0.5 0.0

0.0

0.1

0.2

0.3

0.4

0.5

-1.0 -0.5 0.0

0.0

0.1

0.2

0.3

0.4

0.5

y/S -1.0 -0.5 0.0

0.0

0.1

0.2

0.3

0.4

0.5

z

H

-1.0 -0.5 0.0

0.0

0.1

0.2

0.3

0.4

0.5

-1.0 -0.5 0.0

0.0

0.1

0.2

0.3

0.4

0.5

y/S -1.0 -0.5 0.0

0.0

0.1

0.2

0.3

0.4

0.5

z

H

-1.0 -0.5 0.0

0.0

0.1

0.2

0.3

0.4

0.5

-1.0 -0.5 0.0 0.0

0.1

0.2

0.3

0.4

0.5

0

0

-2

0

0

2

2

2

-1.0 -0.5 0.0 y/S 0.0

0.1

0.2

0.3

0.4

0.5 beta: -6 -4 -2 0 2 4 6 8 10 12 14 16 18 20 22 24 26

 [deg]

z

H

-1.0 -0.5 0.0 0.0

0.1

0.2

0.3

0.4

0.5 -50

-50 -54

-46

-54

-54 -54 -58

-1.0 -0.5 0.0 0.0

0.1

0.2

0.3

0.4

0.5 alpha: -78 -74 -70 -66 -62 -58 -54 -50 -46 -42 -38 -34 -30 -26 -22

 [deg]

(a) Total Pressure Coefficient and Secondary Flow Vectors 

(b) Pitch Angle, β [deg]

(c) Yaw Angle, γ [deg]
NASA/TM—2014-218137 11 

Figure 10.—Pitchwise integrations for i = +5.8°. Figure 11.—Pitchwise integrations for i = −36.7°. 

Pitchwise Integrations—Pitchwise integrations of the total-pressure, pitch angle, and yaw angle were 

calculated at each spanwise immersion for i = +5.8° (Fig. 10) and i = −36.7° (Fig. 11). For each angle, the 

high Tu results are compared to the low Tu results. The area-averaged total-pressure coefficients for the 

high Tu data (Fig. 10(a) and (b), green lines) show a reduction in total-pressure variation along most of 

the blade span. In Figure 10(c) and (d), the maxima at z/H = 0.33 of the low Tu results reflects the core of 

the horseshoe vortex (Fig. 8(a)). With a thinner endwall boundary layer in the grid configuration (high 

Tu), the core of the horseshoe vortex, which rises over the transported endwall flow, adopts a lower 

spanwise position of z/H = 0.24 (also see Fig. 7(a)). 

Pitchwise integrations for i = −36.7° are shown in Figure 11. The total-pressure coefficients 

(Fig. 11(a) and (b)) at this incidence angle reflect the boundary layer that remains intact as it passes 

through the blade. As seen in Table 1, the boundary layer is reduced in size by half for the high Tu data. 

Figure 11(b) shows the differences in boundary layer thickness. The total pressure coefficient at midspan 

for the high Tu configuration is comparable to that of the low Tu, which will be shown to exhibit a 

suction-side laminar separation/transition/reattachment. Consistent with the contours shown in Figures 6 

and 8, the pitch and yaw angles (Fig. 11(c) and (d)) do not show any variation due to an increase in Tu. 

Effect of Turbulence on the Blade Loading 

While the exit flowfield is strongly influenced by the inlet boundary layer thickness, the midspan 

loading diagrams more directly reflect the impact of turbulence level on transition and separation/ 

reattachment. The effects of incidence and Reynolds number on the blade loading are show in Figure 12 

for low Tu and Figure 13 for high Tu. All data were acquired at the nominal design exit Mach number 

except as noted, due to facility limitations. Each figure is arranged with the highest Reynolds number 

(4·Reb) on the left and the baseline Reynolds number (Reb) on the right. The angles presented cover the 

entire tested incidence range. 

0.0 0.2 0.4 0.6 0.8

0.0

0.1

0.2

0.3

0.4

0.5

4 5-hole high

5 5-hole high

4 5-hole low

4 3-hole low

5 5-hole low

5 3-hole low

z

H

psg. probe Tu

  in

(a) Cpt (Area Averaged)

0.00 0.05 0.10 0.15 0.20 0.25

0.0

0.1

0.2

0.3

0.4

0.5

z

H

(b) Cpt (Area Averaged)

(detailed view)

0.0 0.2 0.4 0.6 0.8

0.0

0.1

0.2

0.3

0.4

0.5

4 5-hole high

5 5-hole high

4 5-hole low

4 3-hole low

5 5-hole low

5 3-hole low

z

H

psg. probe Tu

  in

(a) Cpt (Area Averaged)

0.00 0.05 0.10 0.15 0.20 0.25

0.0

0.1

0.2

0.3

0.4

0.5

z

H

(b) Cpt (Area Averaged)

(detailed view)

P.S.

exit

metal

angle

S.S.

exit

metal

angle

avg.

exit

metal

angle

-65 -60 -55 -50 -45

0.0

0.1

0.2

0.3

0.4

0.5

z

H

(c)  [deg] (Mass Averaged)

-1.0 0.0 1.0 2.0

0.0

0.1

0.2

0.3

0.4

0.5

z

H

(d)  [deg] (Mass Averaged)

P.S.

exit

metal

angle

S.S.

exit

metal

angle

avg.

exit

metal

angle

-65 -60 -55 -50 -45

0.0

0.1

0.2

0.3

0.4

0.5

z

H

(c)  [deg] (Mass Averaged)

-1.0 0.0 1.0 2.0

0.0

0.1

0.2

0.3

0.4

0.5

z

H

(d)  [deg] (Mass Averaged)
NASA/TM—2014-218137 12 

Figure 12.—Blade loading (low Tu)—effects of incidence and Reynolds at 

number at nominal design exit Mach number. 

sonic

x/Cx 0.0 0.2 0.4 0.6 0.8 1.0 -1.2

-1.0

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Cps

a) i=15.8, 4Reb, M2,i=0.72

sonic

x/Cx 0.0 0.2 0.4 0.6 0.8 1.0 -1.2

-1.0

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Cps

f) i=15.8, 1Reb, M2,i=0.72

sonic

x/Cx 0.0 0.2 0.4 0.6 0.8 1.0 -1.2

-1.0

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Cps

b) i=0.0, 4Reb, M2,i=0.72

sonic

x/Cx 0.0 0.2 0.4 0.6 0.8 1.0 -1.2

-1.0

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Cps

g) i=0.0, 1Reb, M2,i=0.71

sonic

x/Cx 0.0 0.2 0.4 0.6 0.8 1.0 -1.2

-1.0

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Cps

c) i=16.1, 4Reb, M2,i=0.72

sonic

x/Cx 0.0 0.2 0.4 0.6 0.8 1.0 -1.2

-1.0

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Cps

h) i=16.1, 1Reb, M2,i=0.72

sonic

x/Cx 0.0 0.2 0.4 0.6 0.8 1.0 -1.2

-1.0

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Cps

d) i=36.7, 4Reb, M2,i=0.73 sonic

x/Cx 0.0 0.2 0.4 0.6 0.8 1.0 -1.2

-1.0

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Cps

i) i=36.7, 1Reb, M2,i=0.67

sonic

x/Cx 0.0 0.2 0.4 0.6 0.8 1.0 -1.2

-1.0

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Cps

e) i=51.0, 4Reb, M2,i=0.72

sonic

x/Cx 0.0 0.2 0.4 0.6 0.8 1.0 -1.2

-1.0

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Cps

j) i=51.0, 1Reb, M2,i=0.62

blade %span

6 50%

6 10%

4 50%

4 30%

4 15%

4 10%

5 50%

5 30%

5 15%

5 10%
NASA/TM—2014-218137 13 

Figure 13.—Blade loading (high Tu)—effects of incidence and Reynolds at 

number at nominal design exit Mach number. 

sonic

x/Cx 0.0 0.2 0.4 0.6 0.8 1.0 -1.2

-1.0

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Cps

a) i=15.8, 4Reb, M2,i=0.72

sonic

x/Cx 0.0 0.2 0.4 0.6 0.8 1.0 -1.2

-1.0

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Cps

f) i=15.8, 1Reb, M2,i=0.72

sonic

x/Cx 0.0 0.2 0.4 0.6 0.8 1.0 -1.2

-1.0

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Cps

b) i=0.0, 4Reb, M2,i=0.72

sonic

x/Cx 0.0 0.2 0.4 0.6 0.8 1.0 -1.2

-1.0

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Cps

g) i=0.0, 1Reb, M2,i=0.72

sonic

x/Cx 0.0 0.2 0.4 0.6 0.8 1.0 -1.2

-1.0

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Cps

c) i=16.1, 4Reb, M2,i=0.72

sonic

x/Cx 0.0 0.2 0.4 0.6 0.8 1.0 -1.2

-1.0

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Cps

h) i=16.1, 1Reb, M2,i=0.71

sonic

x/Cx 0.0 0.2 0.4 0.6 0.8 1.0 -1.2

-1.0

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Cps

d) i=36.7, 4Reb, M2,i=0.72 sonic

x/Cx 0.0 0.2 0.4 0.6 0.8 1.0 -1.2

-1.0

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Cps

i) i=36.7, 1Reb, M2,i=0.69

sonic

x/Cx 0.0 0.2 0.4 0.6 0.8 1.0 -1.2

-1.0

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Cps

e) i=51.0, 4Reb, M2,i=0.72

sonic

x/Cx 0.0 0.2 0.4 0.6 0.8 1.0 -1.2

-1.0

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Cps

j) i=51.0, 1Reb, M2,i=0.61

blade %span

6 50%

6 10%

4 50%

4 30%

4 15%

4 10%

5 50%

5 30%

5 15%

5 10%
NASA/TM—2014-218137 14 

At the highest positive incidence and the baseline Reynolds number (Reb) condition, the low Tu flow 

exhibits suction-surface separation and reattachment at x/Cx = 0.85. This is demonstrated as a neutral 

pressure gradient region, followed by an abrupt diffusion (Ref. 14). The high Tu blade loading in 

Figure 13(a) and (f) show that the flow is attached over the entire blade surface—results are comparable 

to the low Tu case at the higher 4·Reb flow condition. For both Tu cases, the positive incidence results in 

high blade loading. The three-dimensionality of the flow can be seen in the spanwise loading variations in 

Figure 12(a) and (f). The magnitude of variations is reduced at high Tu and periodicity is improved. Note 

that for both Tu levels, small supersonic overspeed flow regions exist on the suction surface near the 

leading edge. 

As seen in the total pressure coefficient contours of Figure 6, as the incidence decreases the blade 

unloads and the flow becomes more two-dimensional. For both Tu cases, negative loading is measured on 

the front portion of the blade for i  –16.1°. At i = –36.7° pressure-side cove separation is observed at low 

Tu but the flow remains attached at high Tu. In fact, at high Tu, neither suction-side nor pressure-side 

separation were observed for any operating point within the full range of text-matrix incidence and 

Reynolds number. The effects of a thinner boundary layer and increased Tu reflect that the flow over the 

blade becomes two-dimensional very quickly with decreasing incidence. The only time the blade 

indicates signs of secondary flows (spanwise variation) is at the most positive incidence angle. 

The effects of exit Mach number variation at fixed Reynolds number on blade loading was 

investigated. The results from the low Tu study showed that as the exit Mach number is reduced from the 

nominal M2,i = 0.72 to M2,i = 0.35, the loading increases and the location of minimum Cps moves forward 

(Ref. 4). The same results were found when tested at high inlet Tu; therefore those results will not be 

presented here. 

Midspan Exit Total-Pressure and Flow Angle Surveys 

Midspan total-pressure and exit flow angle surveys were acquired for each incidence angle. Each 

survey consisted of 132 points concentrated in the wakes of three passages. The effects of Tu, Reynolds 

number, and Mach number variations on the wake profile are presented in Figures 14 to 18. At the high 

positive incidence angle of i = +10.8° (Fig. 14), it is noted that the wake width and depth increases with 

decreasing Reynolds number and Mach number. At a high positive incidence the blade row is highly 

loaded and the large suction side thickening is consistent with the separation (low Tu only) and high 

loading observed in the blade loading diagrams. For the high Tu case (Fig. 14(b)), there is minimal 

thickening of the suction-side boundary-layer/wake as compared to the same in the low inlet Tu case 

(Fig. 14(a)). It is noted that wakes are more periodic for the high Tu and thinner inlet boundary layer. 

Both Tu cases show that the exit flow angles are independent of Reynolds number and decrease as the 

Mach number is reduced. 

(a) Low Tu (b) High Tu

Figure 14.—Effects of Reynolds number and exit Mach number at i = +10.8°. 

Passage 4 Passage 5

PS exit metal angle

avg exit metal angle

SS exit metal angle

-2.0 -1.5 -1.0 y/S -0.5 0.0 0.5 -65

-60

-55

-50

-45

-40

-35 2

[deg]

PS SS PS SS PS SS

-2.0 -1.5 -1.0 y/S -0.5 0.0 0.5 -0.1

0.0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8 4.0des 0.72

2.0des 0.72

1.0des 0.72

1.0des 0.35

0.4des 0.35

Pt,1  Pt

Pt,1  P2

Cpt

Cpt =

4.0Reb 0.72

2.0Reb 0.72

1.0Reb 0.72

1.0Reb 0.35

0.4Reb 0.35

ReCx,2 Ma2,i

Passage 4 Passage 5

PS exit metal angle

avg exit metal angle

SS exit metal angle

-2.0 -1.5 -1.0 y/S -0.5 0.0 0.5 -65

-60

-55

-50

-45

-40

-35 2

[deg]

PS SS PS SS PS SS

-2.0 -1.5 -1.0 y/S -0.5 0.0 0.5 -0.1

0.0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8 4.0des 0.72

2.0des 0.72

1.0des 0.72

1.0des 0.35

0.4des 0.35

Pt,1  Pt

Pt,1  P2

Cpt

Cpt =

4.0Reb 0.72 2.0Reb 0.72 1.0Reb 0.72 1.0Reb 0.35 0.4Reb 0.35

ReCx,2 Ma2,i
NASA/TM—2014-218137 15 

(a) Low Tu (b) High Tu

Figure 15.—Effects of Reynolds number and exit Mach number at i = 0.0°. 

(a) Low Tu (b) High Tu

Figure 16.—Effects of Reynolds number and exit Mach number at i = –16.1°. 

As the incidence angle decreases, the trends in the total-pressure and exit flow angles remain the same 

but the low Tu exhibits a larger suction side boundary-layer/wake thickening while the high Tu total 

pressure and exit flow angles vary only slightly with decreasing Reynolds number, see Figure 15. As 

incidence decreases to a negative value (Fig. 16(a)), the high Tu wake is very similar to the i = –36.7° wake 

profile at low Tu (Fig. 17(a)), which corresponds to the incidence with minimum loss at low Tu (Ref. 4). 

The highest negative incidence, i = –51.0°, is shown in Figure 18. At low inlet Tu, the pressure side 

separation spans across a majority of the passage (Fig. 18(a)). This massive separation is associated with 

aerodynamic blockage. The PS wake depth and width increase significantly with decreasing Reynolds 

number. The suction side wake width also increases slightly with decreasing Reynolds number. The exit 

flow angles show a significant decrease, near the average exit metal angle, compared to the previous 

incidence angles. The angles vary little with decreasing Reynolds and Mach number except at the highest 

Reynolds number condition (4·Reb). In the previous study (Ref. 4) it was concluded that at this extreme 

negative incidence, as the Reynolds number decreases the PS induced wake thickens substantially causing 

an aerodynamic blockage that resets the aerodynamic throat upstream and effects increased turning 

resulting in an increased negative discharge angle. At the high inlet Tu (Fig. 18(b)), the flow remains 

largely attached and a gross pressure-side separation is not seen. As the Reynolds number decreases, the 

wake thickness increases slightly on the pressure side. At the lowest Reynolds number (0.4·Reb) the wake 

does indicate a pressure-side separation. The exit flow angles remain consistent and vary slightly with 

decreasing Reynolds number. 

Passage 4 Passage 5

PS SS PS SS PS SS

-2.0 -1.5 -1.0 y/S -0.5 0.0 0.5 -0.1

0.0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

4.0des 0.72

2.0des 0.72

1.0des 0.72

1.0des 0.35

0.4des 0.35

Pt,1  Pt

Pt,1  P2

Cpt

Cpt =

ReCx,2 Ma2,i

4.0Reb 0.72 2.0Reb 0.72

1.0Reb 0.72 1.0Reb 0.35 0.4Reb 0.35

PS exit metal angle

avg exit metal angle

SS exit metal angle

-2.0 -1.5 -1.0 y/S -0.5 0.0 0.5 -65

-60

-55

-50

-45

-40

-35 2

[deg]

Passage 4 Passage 5

PS SS PS SS PS SS

-2.0 -1.5 -1.0 y/S -0.5 0.0 0.5 -0.1

0.0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

4.0des 0.72

2.0des 0.72

1.0des 0.72

1.0des 0.35

0.4des 0.35

Pt,1  Pt

Pt,1  P2

Cpt

Cpt =

ReCx,2 Ma2,i

4.0Reb 0.72 2.0Reb 0.72 1.0Reb 0.72

1.0Reb 0.35 0.4Reb 0.35

PS exit metal angle

avg exit metal angle

SS exit metal angle

-2.0 -1.5 -1.0 y/S -0.5 0.0 0.5 -65

-60

-55

-50

-45

-40

-35 2

[deg]

Passage 4 Passage 5

PS SS PS SS PS SS

-2.0 -1.5 -1.0 y/S -0.5 0.0 0.5 -0.1

0.0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

4.0des 0.72

2.0des 0.72

1.0des 0.72

1.0des 0.35

0.4des 0.35

Pt,1  Pt

Pt,1  P2

Cpt

Cpt =

ReCx,2 Ma2,i

4.0Reb 0.72

2.0Reb 0.72

1.0Reb 0.72

1.0Reb 0.35

0.4Reb 0.35

PS exit metal angle

avg exit metal angle

SS exit metal angle

-2.0 -1.5 -1.0 y/S -0.5 0.0 0.5 -65

-60

-55

-50

-45

-40

-35 2

[deg]

Passage 4 Passage 5

PS SS PS SS PS SS

-2.0 -1.5 -1.0 y/S -0.5 0.0 0.5 -0.1

0.0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

4.0des 0.72

2.0des 0.72

1.0des 0.72

1.0des 0.35

0.4des 0.35

Pt,1  Pt

Pt,1  P2

Cpt

Cpt =

ReCx,2 Ma2,i

4.0Reb 0.72 2.0Reb 0.72

1.0Reb 0.72

1.0Reb 0.35

0.4Reb 0.35

PS exit metal angle

avg exit metal angle

SS exit metal angle

-2.0 -1.5 -1.0 y/S -0.5 0.0 0.5 -65

-60

-55

-50

-45

-40

-35 2

[deg]
NASA/TM—2014-218137 16 

(a) Low Tu (b) High Tu

Figure 17.—Effects of Reynolds number and exit Mach number at i = –36.7°. 

(a) Low Tu (b) High Tu

Figure 18.—Effects of Reynolds number and exit Mach number at i = –51.0°. 

The effects of incidence angle variation at the highest and lowest Reynolds number conditions for the 

low (a) and high (b) Tu cases are summarized in Figures 19 and 20. The effects of an order of magnitude 

variation in Reynolds number are displayed across these figures. The three incidence angles represent the 

mission points; cruise (i = +5.8°), takeoff ( i = –36.7°), and maximum mission incidence (i = –46.0°). For 

the highest Reynolds number (4·Reb) in Figure 19 for both low (a) and high (b) Tu, the positive incidence 

produces the highest overall loss and decreases with decreasing incidence. High Tu produces much lower 

loss than low Tu and varies little at the negative incidence angles. The largest variance between Tu levels 

occur at the low Reynolds number (0.4·Reb) and Mach number (0.35) condition, shown in Figure 20. At low 

Tu (Fig. 20(a)) strong suction side separation is evident at the positive incidence angle. As the incidence 

decreases, the blade unloads (as observed in the blade loading diagrams) and the losses decrease. At the high 

negative incidence (i = –46°) the losses increase due to an extensive pressure-side separation. At high Tu

(Fig. 20(b)) case the wake profile and loading indicate the flow remains attached and the losses decrease. 

The losses are similar to the high Reynolds number, where the loss is the highest at positive incidence and 

decreases with decreasing incidence. At the highest negative incidence angle there is evidence of a slight 

increase on the pressure side. 

For all of the high Tu positive incidence cases, small regions of seemingly erroneous negative Cpt values 

were observed on the PS sides of the wakes (Figs. 14(b), 19(b), and 20(b)). These values were not observed 

for the corresponding low Tu cases; are consistent and repeatable; and are outside the uncertainty range of 

the measurements. Full mapping of the inlet total pressure field was performed for the i = +5.8° case and no 

sufficiently high local Pt,1 values were found. Potential other explanations, including flow unsteadiness 

Passage 4 Passage 5

PS SS PS SS PS SS

-2.0 -1.5 -1.0 y/S -0.5 0.0 0.5 -0.1

0.0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

4.0des 0.73

2.0des 0.72

1.0des 0.72

1.0des 0.35

0.4des 0.35

Pt,1  Pt

Pt,1  P2

Cpt

Cpt =

ReCx,2 Ma2,i

4.0Reb 0.73

2.0Reb 0.72

1.0Reb 0.72

1.0Reb 0.35

0.4Reb 0.35

PS exit metal angle

avg exit metal angle

SS exit metal angle

-2.0 -1.5 -1.0 y/S -0.5 0.0 0.5 -65

-60

-55

-50

-45

-40

-35 2

[deg]

Passage 4 Passage 5

PS SS PS SS PS SS

-2.0 -1.5 -1.0 y/S -0.5 0.0 0.5 -0.1

0.0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

4.0des 0.73

2.0des 0.72

1.0des 0.72

1.0des 0.35

0.4des 0.35

Pt,1  Pt

Pt,1  P2

Cpt

Cpt =

ReCx,2 Ma2,i

4.0Reb 0.73

2.0Reb 0.72

1.0Reb 0.72 1.0Reb 0.35

0.4Reb 0.35

PS exit metal angle

avg exit metal angle

SS exit metal angle

-2.0 -1.5 -1.0 y/S -0.5 0.0 0.5 -65

-60

-55

-50

-45

-40

-35 2

[deg]

Passage 4 Passage 5

PS exit metal angle

avg exit metal angle

SS exit metal angle

-2.0 -1.5 -1.0 y/S -0.5 0.0 0.5 -65

-60

-55

-50

-45

-40

-35 2

[deg]

PS SS PS PS SS SS

-2.0 -1.5 -1.0 y/S -0.5 0.0 0.5 -0.1

0.0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

4.0des 0.72

2.0des 0.72

1.0des 0.62

1.0des 0.35

0.4des 0.35

Pt,1  Pt

Pt,1  P2

Cpt

Cpt =

ReCx,2 Ma2,i

4.0Reb 0.72

2.0Reb 0.72 1.0Reb 0.62

1.0Reb 0.35 0.4Reb 0.35

Passage 4 Passage 5

PS exit metal angle

avg exit metal angle

SS exit metal angle

-2.0 -1.5 -1.0 y/S -0.5 0.0 0.5 -65

-60

-55

-50

-45

-40

-35 2

[deg]

PS SS PS PS SS SS

-2.0 -1.5 -1.0 y/S -0.5 0.0 0.5 -0.1

0.0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

4.0des 0.72

2.0des 0.72

1.0des 0.62

1.0des 0.35

0.4des 0.35

Pt,1  Pt

Pt,1  P2

Cpt

Cpt =

ReCx,2 Ma2,i

4.0Reb 0.72 2.0Reb 0.71

1.0Reb 0.60 1.0Reb 0.35

0.4Reb 0.35
NASA/TM—2014-218137 17 

(a) Low Tu (b) High Tu

Figure 19.—Effect of incidence angle at ReCx,2 = 2.12106

 (4·Reb) and M2,i = 0.72. 

(a) Low Tu (b) High Tu

Figure 20.—Effect of incidence angle at ReCx,2 = 2.12105 (0.4·Reb) and M2,i = 0.347. 

(grid shedding); differing responses to high turbulence for Kiel versus five-hole probes; and the impact of 

streamwise velocity gradients on the local turbulent kinetic energy (production/destruction), were 

considered, but no viable justification could be found. Consideration of this repeatable, and yet apparently 

anomalous, total-pressure gradient continues. 

Loss Bucket and Flow Turning 

For both Tu cases, area-averaged integrations of the total-pressure data were used to calculate the 

midspan profile loss coefficients, ω, for all 10 incidence angles and five flow conditions. The loss buckets 

(i.e., the midspan loss coefficients plotted as a function of incidence) for low and high Tu are provided in 

Figures 21(a) and (b). The integrations were calculated separately over two complete passages, 4 and 5. 

At low Tu and extreme negative incidence angles, passage-to-passage variations are quite evident in the 

scatter of the loss levels (Fig. 21(a)). These variations are attributable to the sensitivity of transitional flow 

with separation. Periodicity is improved at high Tu, reducing scatter in the loss levels from passage to 

passage (Fig. 21(b)). The sensitivity of loss levels to Reynolds number at high Tu is greatly reduced and 

the losses collapse asymptotically to the high Reynolds number (turbulent) level. Loss levels remain 

nearly constant as the Mach number is varied from M2,i = 0.72 to M2,i = 0.35 at 1·Reb. It is interesting to 

note that the apparent pinch-point (invariance of loss with Re) at i = –36.7° of the low Tu data noted 

earlier (Ref. 4), was exhibited in high Tu data, arguably to an even greater degree. 

 

PS SS PS SS PS SS

y/S -2.0 -1.5 -1.0 -0.5 0.0 0.5 -0.1

0.0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

i = 5.8

i = 36.7

i = 46.0

Cpt Pt,1  Pt Cpt = 

Pt,1  P2

Passage 4 Passage 5

PS SS PS SS PS SS

y/S -2.0 -1.5 -1.0 -0.5 0.0 0.5 -0.1

0.0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

i = 5.8

i = 36.7

i = 46.0

Cpt Pt,1  Pt Cpt = 

Pt,1  P2

Passage 4 Passage 5

PS SS PS SS PS SS

y/S -2.0 -1.5 -1.0 -0.5 0.0 0.5 -0.1

0.0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

i = 5.8

i = 36.7

i = 46.0

Pt,1  Pt Cpt = 

Pt,1  P2

Cpt

Passage 4 Passage 5

PS SS PS SS PS SS

y/S -2.0 -1.5 -1.0 -0.5 0.0 0.5 -0.1

0.0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

0.8

i = 5.8

i = 36.7

i = 46.0

Pt,1  Pt Cpt =  Pt,1  P2

Cpt

Passage 4 Passage 5
NASA/TM—2014-218137 18 

(a) Low Tu (b) High Tu

Figure 21.—Midspan loss versus incidence. 

(a) Low Tu (b) High Tu

Figure 22.—Re–n scaled midspan loss versus incidence. 

The corrected midspan loss levels are plotted as a function of a corrected Reynolds number in 

Figures 22(a) and (b). The low Tu data strongly collapsed on the power-law ωRe–0.5, (Fig. 22(a)), 

indicative of a significant influence of laminar flow on the midspan loss levels. For the high Tu data 

(Fig. 22(b)), a strong collapse was achieved with ωRe-0.1, indicating that, as expected, the midspan 

losses were dominantly influenced by turbulent flow. 

As carried out in McVetta et al. (Ref. 4), the high and low Tu midspan loss coefficients were 

re-plotted in terms of reduced loss and reduced incidence according to the Ainley-Mathieson (A-M) 

scaling (ω/ωs versus i/is) (Ref. 15) as shown in Figure 23. The strong collapse found previously for low Tu

data (Ref. 4) was repeated, as expected, for the high Tu data. It is noted that, if A-M scaling is applied 

about iopt = –26°, as shown in Figure 24, then not only do the data collapse strongly on the A-M scaling, 

but the A-M loss correlation (Ref. 15) was found to hold. With respect to this new reference zero 

incidence and associated loss level, the A-M correlation might be used to extrapolate negative incidence 

angles out to a range of angles that is far beyond the range of the present experiment. 

The midspan deviation angle from the exit metal angle as a function of incidence and Reynolds 

number is shown in Figure 25(a) (low Tu) and Figure 25(b) (high Tu). At low Tu, the deviation angle 

asymptotically approaches Δβ2 = 2° as the incidence decreases at the design exit Mach numbers and 

approaches Δβ2 = 4° at the lower Mach condition. As the angle approaches the extreme negative 

incidence angle, where the aerodynamic blockage on the pressure side causes a more negative discharge 

 

0

0.05

0.1

0.15

0.2

0.25

0.3

-60 -50 -40 -30 -20 -10 0 10 20

ω (Area Averaged)

i, Incidence [deg]

4        0.72        4

2        0.72        4

1        0.72        4

1        0.35        4

0.4     0.35        4

ሺ࢞ࢉࢋࡾ,૛ሻ/〖ࢋࡾ_〖࢈ M2,i

4.0 0.72 4 Passage

4.0 0.72 5

2.0 0.72 4

2.0 0.72 5

1.0 0.72 4

1.0 0.72 5

1.0 0.35 4

1.0 0.35 5

0.4 0.35 4

0.4 0.35 5

Re

Cx,2

Reb M2,i passage

0

0.05

0.1

0.15

0.2

0.25

0.3

0.35

0.4

-60 -50 -40 -30 -20 -10 0 10 20

ωcorrected (Area Averaged)

i, Incidence [deg]

4        0.72        4

4        0.72        5

2        0.72        4

2        0.72        5

1        0.72        4

1        0.72        5

1        0.35        4

1        0.35        5

0.4        0.35        4

0.4        0.35        5

࢈_〖ࢋࡾ〗/ሻ૛,࢞ࢉࢋࡾሺ

4.0 0.72 4

4.0 0.72 5

2.0 0.72 4

2.0 0.72 5

1.0 0.72 4

1.0 0.72 5

1.0 0.35 4

1.0 0.35 5

0.4 0.35 4

0.4 0.35 5

Re

Cx,2

Reb M2,i passage
NASA/TM—2014-218137 19 

angle, the deviation angle shifts significantly. At high Tu, the flow remains attached on the blade surface 

for the entire incidence range and the deviation angle asymptotically approaches 2°  Δβ2  3° for all 

Reynolds numbers; differences (offset) between the baseline (0.72) and lower (0.35) exit Mach number 

conditions were again evident (Fig. 25(b)). 

 (a) Low Tu (b) High Tu

Figure 23.—Midspan loss bucket on ainley-mathieson scaling. 

Figure 24.—Midspan loss bucket on ainley-mathieson scaling at high Tu. 

 (a) Low Tu (b) High Tu

Figure 25.—Midspan deviation angle from exit metal angle as function of incidence and Reynolds number. 

0

1

2

3

4

5

6

7

8

9

-15 -10 -5 0 5

ω /i = 0

i / is

4        0.72        4

2        0.72        4

1        0.72        4

1        0.35        4

0.4     0.35        4

M2,i Passage

0

1

2

3

4

5

6

7

8

9

-4 -3 -2 -1 0 1 2

ω /i = iopt

(i - iopt) / (is - iopt)

4        0.72        4

2        0.72        4

1        0.72        4

1        0.35        4

0.4     0.35        4

Ainley‐Mathieson correlation

M2,i Passage

-2

0

2

4

6

8

10

12

14

-60 -50 -40 -30 -20 -10 0 10 20

Δβ2 [deg]

i, Incidence [deg]

4        0.72        4

4        0.72        5

2        0.72        4

2        0.72        5

1        0.72        4

1        0.72        5

1        0.35        4

1        0.35        5

1        0.35        4

1        0.35        5

ሺ࢞ࢉࢋࡾ,૛ሻ/〖ࢋࡾ_〖࢈ M2,i

4.0 0.72 4

4.0 0.72 5

2.0 0.72 4

2.0 0.72 5

1.0 0.72 4

1.0 0.72 5

1.0 0.35 4

1.0 0.35 5

0.4 0.35 4

0.4 0.35 5

Re

Cx,2

Reb M2,i passage

‐2

0

2

4

6

8

10

12

14

‐60 ‐50 ‐40 ‐30 ‐20 ‐10 0 10 20

Δβ2 [deg]

i, Incidence [deg]

4.0   0.72   4

4.0   0.72   4

2.0   0.72   4

2.0   0.72   5

1.0   0.72   4

1.0   0.72   5

1.0   0.35   4

1.0   0.35   5

0.4   0.35   4

0.4   0.35   5

M2,i Passage

4.0 0.72 4

4.0 0.72 5

2.0 0.72 4

2.0 0.72 5

1.0 0.72 4

1.0 0.72 5

1.0 0.35 4

1.0 0.35 5

0.4 0.35 4

0.4 0.35 5

Re

Cx,2

Reb M2,i passage
NASA/TM—2014-218137 20 

Conclusions 

A new set of aerodynamic performance data was obtained for VSPT blading at engine relevant 

(high Tu) inlet turbulence conditions. The dataset complements an earlier dataset acquired at low inlet 

turbulence levels as required to achieve mission-relevant transitional flow behavior within the Reynolds 

number capability of the experimental facility. The high Tu data were acquired over the same wide range 

of incidence angles and repeated the Mach/Reynolds number conditions of the earlier low Tu test matrix. 

As in the earlier low Tu work, exit total-pressure and exit flow angles were examined to understand 

the strength and impact of secondary flow fields at the highly loaded design incidence (i = +5.8°) 

condition of LCTR cruise and at the low-load, high negative incidence (i = –36.7°) of LCTR takeoff. The 

turbulence grid of the high Tu entry effectively reduced by half the thickness of the turbulent endwall 

boundary layers at cascade inlet, leading to significantly less aerodynamic blockage in the test section. 

Comparison of blade loading diagrams at the low Tu and high Tu conditions showed that the flow at 

the high Tu conditions remains attached on both SS and PS at all incidence, Reynolds number, and 

Mach number conditions. This is in strong contrast to the blade loading at low Tu which exhibited 

separation, transition, reattachment on both suction and pressure sides of the blade, depending on 

incidence angle. The pressure side cove separation, a key feature in the data at low Tu, was not evident in 

the high Tu data. Given that the lower mission-relevant Reynolds numbers were not achievable in the 

NASA turbine cascade test facility, it remains for separate work (Refs. 16 and 17) to assess transitional 

flow characteristics at combined high Tu and cruise-relevant Reynolds numbers. 

Midspan loss levels at the highly loaded design incidence at high Tu were found to be slightly lower 

than at low Tu where the flow separated and reattached. The midspan loss at high Tu was found to be 

higher than at low Tu at lower load conditions at which the low Tu flow did not separate on the pressure 

or suction sides. Spanwise profiles of loss indicated minimum turning was achieved at the spanwise 

location at which the pressure-side leg of the horseshoe vortex exited the blade row. This position change 

in the two entries (low and high Tu) was attributable to the magnitude of the incoming turbulent endwall 

boundary-layer with and without the grid in place and the subsequent impact on the amount of 

low-momentum flow (aerodynamic blockage) at the hub/suction-side corner. 

Midspan profile loss buckets at high Tu collapsed strongly on a turbulent power-law scaling 

(  Re–0.1) reflective of turbulent flow conditions. As in previous low Tu work, Ainley Mathieson 

scaling illustrated strong data collapse. Unlike at low Tu, the collapse was found to be independent of 

Reynolds number and, given proper selection of reference incidence, in strong agreement with the 

Ainley-Mathieson profile loss versus incidence correlation. 

References 

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

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