Jesse Christophel

Pratt & Whitney, East Hartford, Connecticut

Understanding and Mitigating Tip Leakage and

Endwall Losses in High Pressure Ratio Cores

NASA/CR—2015-218854

August 2015
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Jesse Christophel

Pratt & Whitney, East Hartford, Connecticut

Understanding and Mitigating Tip Leakage and

Endwall Losses in High Pressure Ratio Cores

NASA/CR—2015-218854

August 2015

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared under Contract NNC10BA12B, Task order NNC11TA41T
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ABSTRACT 

Reducing endwall and tip secondary flow losses will be a key enabler for the next generation of 

commercial and military air transport and will be an improvement on the state-of-the-art in 

turbine loss reduction strategies. 

The objective of this research is three-fold: 

1) To improve understanding of endwall secondary flow and tip clearance losses 

2) To develop novel technologies to mitigate these losses and test them in low-speed 

cascade and rig environments 

3) To validate predictive tools 

To accomplish these objectives, Pratt & Whitney (P&W) has teamed with Pennsylvania State 

University (PSU) to experimentally test new features designed by P&W. P&W will create new 

rim-cavity features to reduce secondary flow loss and improve purge flow cooling effectiveness 

and new blade tip features to manage leakage flows and reduce tip leakage secondary flow loss. 

P&W is currently developing technologies in these two areas that expect to be assimilated in the 

N+2/N+3 generation of commercial engines. 

 

Understanding and Mitigating Tip Leakage and Endwall Losses in 

High Pressure Ratio Cores 

Jesse Christophel 

Pratt & Whitney 

East Hartford, Connecticut 06108 

NASA/CR—2015-218854 1
INTRODUCTION 

P&W has designed a single stage, high pressure turbine for testing in the Axial Flow Turbine 

Research Facility (AFTRF). This design differs from an engine because the test facility is lowspeed and low-temperature. The AFTRF test turbine has also been scaled up relative to the 

engine to allow for better spatial resolution. 

After designing the airfoils for the AFTRF, additional changes were made to suit the geometry to 

the Low Speed Cascade (LSC). These changes include scaling up by 3X (relative to the AFTRF) 

to fit the cascade dimensions, transforming the 3D vane into a 2D geometry, and replacing the 

blade with an elliptical bluff body that will represent the pressure distortion imposed on the rim 

cavity flow by a blade. 

The purpose of this final technical report is to summarize key findings from work completed 

during NRA contract NNC10BA12B, task order NNC11TA41T (August 2011 – December 

2014). This work was completed by the primary contract holder, Pratt & Whitney, and the 

subcontractor, The Pennsylvania State University. Additional test results and details are 

included in the attachments, which contain the university thesis documents written by the 

students performing this contract work. 

 

NASA/CR—2015-218854 2
TECHNICAL DISCUSSION 

1 Facility overview 

Testing under this NRA occured in two separate but complimentary facilities. The first is the 

AFTRF, which is a low-speed rig. This facility allows for steady and unsteady measurements in 

the static and rotating frame of reference. These measurements are critical to understanding the 

unsteady wakes and secondary flow structures that result from blade tip and rim cavity leakage 

flow. A cross-section of the AFTRF is shown in Figure 1. Flow enters through a bell-mouth 

inlet, passes through the single-stage turbine, and then exits through a guide vane. The geometry 

shown in this cross section is from previous testing and does not represent the P&W hardware. 

Figure 1: Axial Flow Turbine Research Facility Cross Section 

The second facility is the LSC, which is a re-circulating wind tunnel with a removable test 

section located at one corner. This cascade is larger scale than the AFTRF which allows for 

greater measurement resolution. Because of the stationary nature of the cascade, it allows for a 

focused study on endwall and rim cavity flow structures. To represent the effect of a passing 

blade in the engine environment, a bluff body will be positioned downstream of the rim cavity. 

This will allow for the simulation of an instantaneous blade clocking position while measuring 

purge effectiveness and characterizing the flow field where the purge flow interacts with the 

mainstream. Because of the relative ease of testing and lower cost associated with the LSC, 

many geometries can be tested, and only the most promising will be installed in the AFTRF for 

rotating tests. An overview of the LSC is shown in Figure 2. 

NASA/CR—2015-218854 3
 Figure 2: Low Speed Cascade Wind Tunnel Overview 

Test Geometries 

A comparison of the AFTRF and LSC cross sections is shown in Figure 3. Note that in this 

image the cross sections have been scaled to the same size. This figure shows the similarity in 

geometry between the AFTRF and LSC. 

Figure 3: AFTRF (solid) and LSC (dashed) Cross Section Overlay 

The rim cavity shown in Figure 3 is a basic overlapping seal where the vane platform trailing 

edge overlaps the blade platform leading edge. The purge flow enters at the bottom through an 

opening created by a typical knife-edge engine seal. In the AFTRF, this seal gap will be 

measured and controlled and in the LSC a simple slot opening will be used to introduce the air. 

The test vane is based on a modern design for high pressure turbine first stage vanes. The LSC 

vane has a constant cross section based on the 25% span cross section of the AFTRF vane. A 

solid line on the AFTRF vane in Figure 4 shows where the LSC vane cross section comes from. 

NASA/CR—2015-218854 4
Figure 4: AFTRF (left) and LSC (right) Vane Geometry 

A summary of the scaling parameters considered and final design values for the test vanes is 

given in Table 1. A scale factor for the LSC of 3.0, relative to the AFTRF geometry, was chosen 

based on the constraints of wind tunnel flow capability and required number of vanes to ensure 

flow periodicity. For both test vanes, the inlet Reynolds number was maintained. The inlet 

velocity, exit velocity, and flow angle are mass averaged values obtained from steady RANS 

CFD of both the LSC and the AFTRF. The flow angle values are at a streamwise location 10% 

axial chord downstream of the vane trailing edge. 

Table 1: Scaling Parameters and Design Summary 

Figure 5 shows the surface static pressure distribution for the two vanes as well as a vane cross 

section at 25% span. The pressure distributions from the LSC and AFTRF were obtained from 

the same steady RANS CFD as the flow parameters listed above in Table 1. These cross sections 

have been scaled to the same axial chord for easier comparison. As the figure shows, there has 

been a minor change to the LSC vane cross section to account for the incompressible flow 

environment. 

NASA/CR—2015-218854 5
Figure 5: Vane Cross Section and Pressure Distribution Comparison 

Additional checks on the vane design were to match the vane wake pressure distortion imposed 

on the rim cavity. This pressure distortion greatly affects the flow near the rim seal exit, which is 

the area of concern for these tests. Figure 6 shows a comparison between the AFTRF and the 

LSC taken at 10% and 25% span on a plane located at the rim cavity leading edge. The figures 

on the right show the axial plane and the various span values from which the vane wake pressure 

distributions were taken. The plot on the left is of local pressure coefficient at various spans. The 

pressure distribution was non-dimensionalized using mass-averaged reference values from the 

plane and spans at which the pressure distribution was taken. This local pressure coefficient 

allows the comparison of the annular AFTRF facility and the linear LSC, which appears very 

similar. Slight differences are seen between the two cases in which the LSC has a slightly 

smaller difference between the maximum and minimum values of the pressure coefficient 

distribution. Figure 6 illustrates that the vane wake pressure distortion at the rim seal exit is very 

similar between the two facilities. 

NASA/CR—2015-218854 6
Figure 6: Test Vane Pressure Distortion at Rim Cavity Edge 

The test blade is based on a modern design for high pressure turbine first stage blades. An image 

of the AFTRF test blade is shown in Figure 7. This blade was designed so that the tip and 

endwall sections are replaceable. This will allow for easy changes of test geometry. 

Figure 7: AFTRF Blade Geometry 

Figure 8shows the surface static pressure distribution for the AFTRF blade as well as a blade 

cross section at 25% span. 

NASA/CR—2015-218854 7
Figure 8: Vane Cross Section and Pressure Distribution 

For the LSC, a bluff body design was chosen from available NACA airfoils to provide the 

correct upstream pressure distortion that the blade imposes on the rim-cavity. A cross-section of 

the bluff body, as installed in the LSC is shown in Figure 9. As you can see, the stationary bluff 

bodies are relatively small compared to the vane or a traditional blade. 

Figure 9: LSC Bluff Bodies 

The pressure distortion from the AFTRF blade and the LSC bluff bodies are compared in Figure 

10. The two plots of pressure are taken at the rim-cavity leading edge and trailing edge and this 

figure shows that the LSC bluff bodies have excellent representation of the upstream pressure 

effect on the rim-cavity due to the downstream blade. 

NASA/CR—2015-218854 8
Figure 10: Rim-Cavity Static Pressure Distortion Due to 

AFTRF Blade and LSC Bluff Body 

P&W and PSU have created a single-stage turbine design for the AFTRF and the LSC that have 

matched aerodynamic characteristics for the two facility operating parameters. The inlet 

Reynolds number and stage pressure ratios are matched, the vane pressure distributions are 

matched, and the rim-cavity static pressure field is matched with the presence of a downstream 

blade or bluff body. 

Based on this, the two experimental facilities have been set-up correctly to perform testing on 

similar geometries. 

NASA/CR—2015-218854 9
2 Low speed cascade 

The test section for the LSC is shown in Figure 11. In this image, you can see the five vane 

leading edges and four central flow paths. The outer flow paths are controlled with flexible 

walls and by-pass tubes to help balance the flow evenly through the four central passages. 

Figure 11: NRA Test Section Used in the LSC

To overcome the static nature of the LSC test and get adequate periodicity in the rim-cavity, a 

cross-flow loop was installed to represent the circumferential flow created by a spinning disk and 

blade in the rotating environment. This cross-flow was metered and controlled to provide 

periodic pressure boundary conditions at the rim-cavity exit and is separate from the rim-cavity 

purge exit flow. Figure 12 shows a cross-section of the LSC rim-cavity with the cross flow inlet 

and exit as well as a photograph of the cross-flow loop. 

Figure 12: LSE Test Section Showing Cross-Flow Loop

The vane endwall and downstream rim cavity geometry were manufactured with expanded 

polystyrene foam, shown in Figure 13. The manufacturing method permitted each component to 

be cut from one piece, resulting in accurate geometry definition. The downside of this material 

NASA/CR—2015-218854 10
is that the surface is much rougher, so a low conductivity paste, called “Temp-Coat 101” was 

applied to the surface to achieve the desired smoothness. 

Figure 13: Downstream Side of Rim Cavity and Endwall 

Made With Expanded Polystyrene Foam 

Instrumentation of LSC Hardware 

Below is a comprehensive list of instrumentation on the LSC: 

1. Boundary condition measurements 

a. Inlet velocity profile (Pitot probe) 

b. Inlet temperature profile (probe-mounted thermocouple) 

c. Vane surface static pressure to ensure periodicity (surface mounted static pressure 

taps) 

d. Bluff body surface static pressure to ensure periodicity (surface mounted static 

pressure taps) 

e. Coolant pressure and temperature (plenum mounted static pressure tap and 

thermocouples) 

f. Coolant flow rate (in-line venturi meter) 

g. Cross flow inlet and exit temperature (to ensure no heat-up or cool-down of cross 

flow during return loop) 

h. Cross flow rate (in-line venture meter) 

2. Aerodynamic losses 

a. Downstream total pressure (5-hole pressure probe) 

3. Rim cavity effectiveness 

a. Endwall adiabatic film effectiveness (IR camera) 

b. Purge flow temperature (rim seal thermocouple rakes) 

c. Seal effectiveness with species tracking (CO2 sniffers) 

4. Main stream flow field 

a. 2D flow field measurements (PIV measurement system) 

The row of vane static pressure taps was built into each vane at the mid span location. These 

taps will be used to provide benchmark information as well as obtain vane to vane pressure field 

periodicity. The static pressure taps were built into the SLA models so that instrumentation 

NASA/CR—2015-218854 11
connections could be more easily made below the test section end wall. The LSC vanes and 

static pressure taps are shown in Figure 14. 

Figure 14: LSC Vane Showing Static Pressure Tap Holes 

Rim cavity measurement locations are shown in Figure 15. Thermocouple rakes and pressure 

taps are be placed in the seal cavity over two vane pitches at an interval of 20% of the vane pitch. 

These measurements will complement the surface IR measurements of adiabatic wall 

temperature on the mainstream endwalls to determine rim cavity flow effectiveness. 

Figure 15: LSC Rim Cavity Pressure and Temperature Measurements. Red rectangle in 

 top down view represents regions if IR camera thermography. 

NASA/CR—2015-218854 12
Aerodynamic loss is measured using a five-hole pressure probe downstream of the vane row. 

The probe and approximate measurement location is shown in Figure 16. The necessary inlet 

total pressure and velocity needed for this loss calculation will be monitored throughout testing 

as it is necessary to confirm uniform, steady state inlet conditions. 

Figure 16: LSC Five-Hole Pressure Probe and Measurement Location 

The rim-cavity CO2 measurement ports are shown in Figure 17. Existing static pressure taps 

were connected to a CO2 sniffer for these measurements. 

Figure 17: LSC Rim-Cavity Static Pressure Tap and CO2 Measurement Locations 

The LDV measurement planes are shown in Figure 18. The LDV system uses two cameras 

(stereo-PIV) to capture all three components of velocity within the plane of measurement. 

NASA/CR—2015-218854 13
Figure 18: LSC Rim-Cavity Stereo PIV Measurement Planes 

The LSC test section has been designed with adequate controls for ensuring periodic flow 

through the vanes, over the rim-cavity, and through the bluff bodies. Main stream flow and 

purge flow are both monitored and controlled and rim-cavity cross-flow is controlled to ensure 

periodic rim-seal boundary conditions. 

 

NASA/CR—2015-218854 14
Low Speed Cascade Results 

Initial testing of the vane cascade involved characterizing the static pressure distribution on the 

vanes at mid-span to ensure that periodic flow occurred in the mainstream. Figure 19 shows 

good periodicity among the vanes in the test section with and without the purge flow. 

Figure 19. Mid-span pressure distribution on vanes with and without the rim seal

Static pressure distributions on the endwall with no purge flow agreed well with that predicted 

from computational fluid dynamics (CFD) from both a periodic solution as well as a five-vane 

solution. Those results are shown in Figure 20 with good agreement between the experiments 

and both CFD solutions over the majority of the passages. The static pressure measurements 

were made from a point just downstream of the vanes and upstream of the seal (labelled tap 1 in 

Figure 17). 

NASA/CR—2015-218854 15
Figure 20. Comparison of vane endwall pressure distribution from the LSC and CFD predictions

In a gas turbine engine, tangential flow is induced in the seals resulting from the swirl in the 

mainstream created by the vanes and the rotation of the downstream rotor. Since the experiments 

conducted in the LSC are static experiments, it was necessary to simulate the cross-flow that 

actually occurs in the operating engine. In addition, initial static pressure measurements on the 

platform in the seal region indicated a strong aperiodic behavior of the flow exiting the seal. The 

aperiodicity was verified using (CFD) simulations replicating the five vane test section with the 

finite number of vanes. Those aperiodic results are shown in Figure 21 where the streamlines are 

colored by the fluid temperature, which is also used to quantify the sealing effectiveness: 

Figure 21 indicates that there is a fluid build-up on the inner wall (top side of figure) which is 

influenced by the external cross-flow and the vane turning. To alleviate the aperiodic behavior 

and more closely simulate that which occurs in a turbine with multiple passages, a cross-flow 

loop was designed through the use of continued CFD studies. 

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Tஶ െ T௖

NASA/CR—2015-218854 16
Figure 21: Streamlines and contours of pressure coefficient from the LSC five vane CFD simulations

Numerous CFD simulations were conducted to guide the modifications made to the LSC test 

section. The objectives of the CFD simulations were to ensure periodic flow in the region where 

the seal interfaces with the mainstream. Resulting from these simulations was the addition of a 

cross-flow loop in the seal exit region as shown in Figure 22. As can be seen from the 

illustrations, a flow loop containing a blower and a laminar flow element (LFE) was constructed 

and installed into the test section. The flow was introduced into the seal exit region at the outer 

vane location and the same flow was removed from the seal exit region at the inner vane 

location. The amount of cross-flow was determined through the CFD studies to meet the 

periodic objectives. 

Figure 22. Overview of cross-flow loop with detail of cross-flow injection site

 

The effect of the cross-flow velocities induced by the flow loop can be seen in Figure 23 

compared with the CFD solution of the infinite periodic cascade for a 0.5% purge flow rate from 

the seal. As the cross flow velocity ratio is increased from none (0.0) to 2.3, the static pressure 

coefficient more closely agrees with the CFD pressure coefficient for the periodic case. Note 

that the cross flow velocity ratio is defined as: 

NASA/CR—2015-218854 17
 

The Vθ value is the mass-averaged velocity in the tangential direction whereas Uin is the vane 

inlet velocity. CFD simulations were conducted for each of the purge flows to determine the 

proper cross flow velocity ratio. The results of those simulations are shown in Table 2. It is 

interesting to note that as the purge flow increases, the cross flow velocity ratio decreases, which 

is expected given the strong purge flow influence. Comparisons of the CFD results were also 

made to the AFTRF in which the blades are rotating. The CFD predictions of the cross flow 

velocity ratio were found to be in good agreement with the AFTRF as shown in Table 3. 

Figure 23. Variation of the vane endwall pressure distribution as a function of cross-flow rate

Table 2: Comparison of Optimal Cross-Flow Rates from the Periodic CFD and those from the LSC

Table 3: Comparison of the Optimal Cross-Flow Rate from the LSC and that from the AFTRF

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NASA/CR—2015-218854 18
As a comparison to the other purge flows, Figure 24 shows the measured platform static pressure 

variations for the cross flow velocity ratios that were determined through the CFD study. As can 

be seen by the data presented, the cross flow velocity ratios gave good periodicity along the 

endwall. 

Figure 24. Vane endwall pressure distribution for various purge rates and cross flow rates

An uncertainty analysis was performed for sealing effectiveness, adiabatic effectiveness, and 

inlet pressure coefficient using the partial differential method described in Moffat (1985). The 

uncertainty of the reported inlet pressure coefficient values ranged from ±0.39 for a low value of 

4.1 to ±0.71 for a high value of 17.4. Uncertainties in adiabatic endwall effectiveness were ±0.02 

for both a low η value of 0.22 and for a high η value of 0.89. Sealing effectiveness was 

calculated with both thermocouples and carbon dioxide concentration sensing. For both methods, 

the uncertainties were dominated by the measurement of the temperature or carbon dioxide of the 

purge air in the plenum. Using the thermocouple method, uncertainties in sealing effectiveness 

were ±0.02 for a low ɂୡ value of 0.50 and a high value of 0.89. The carbon dioxide method 

yielded a similar uncertainty value, ±0.02 for values between ɂୡ = 0.5 and 0.8. Uncertainty of the 

purge flow rates were ±1% at the low value of 0.75% of the mainstream flow rate and ±0.5% at 

the high value of 1.5%. Error bars have been placed on figures where the uncertainty exceeds the 

size of the symbol. 

Uncertainty for the flow field measurements was calculated using the perturbation method by 

Moffat (1985) due to the complex algorithms used to transform the non-orthogonal 

measurements in to an orthogonal coordinate system. Precision error was measured in the shear 

layer where the highest levels of turbulent kinetic energy were observed. This should provide an 

upper bound for the precision error. The uncertainty for the axial velocity, the tangential 

velocity, the radial velocity, and the turbulent kinetic energy were 5.25%, 0.95%, 20.1%, and 

5.4% of their respective average values. 

NASA/CR—2015-218854 19
Sealing Effectiveness with and without the Downstream 

Bluff Body 

Once the facility was properly modified to include the cross-flow loop, sealing effectiveness 

measurements were made for the various purge flow rates with the baseline seal geometry. The 

baseline geometry is shown in Figure 15. Sealing effectiveness measurements were made with 

thermocouple rakes whose pitchwise locations are shown in Figure 15. Each rake consisted of 

five thermocouples and spanned the entire seal exit. Uncertainty in sealing effectiveness was 

calculated to be ±0.02. Results for the various purge flows with the appropriately scaled crossflow velocity ratios are shown in Figure 25. As would be expected, as the purge flow increases, 

the sealing effectiveness also increases.

Figure 25. Rake-averaged sealing effectiveness at the rim seal exit with no bluff bodies present

One of the important questions being addressed in this study is to determine the effect of the 

downstream blade on the sealing effectiveness. As a result of this question, a bluff body was 

designed and constructed for the LSC experiments. The bluff bodies used are shown along with 

the endwall temperature contours in Figure 26 for the case with and without the bluff bodies. 

These temperatures contours give an indication of how quickly the coolant purged from the seal 

dissipates into the freestream. The contours also give an indication of the periodicity of the flow. 

The contours shown in Figure 26 illustrate the effect that the bluff bodies have on the endwall 

cooling pattern as compared to the case with no bluff bodies. In both cases the contours show a 

two-dimensional behavior in the cooling patterns. The results also show that as the purge flow 

rate is increased so is the cooling present on the endwall. 

NASA/CR—2015-218854 20
Figure 26. Comparison of endwall adiabatic effectiveness for the baseline rim seal with and without bluff bodies

The results in Figure 27 show the effect that the bluff bodies have on the overall sealing 

effectiveness for the various purge flows. The overall sealing effectiveness was calculated by 

averaging sealing effectiveness data from all of the thermocouple rakes into a single value. The 

bluff bodies decrease the overall sealing effectiveness. This effect is more pronounced for the 

lower purge flow rates. The largest decrease is approximately 10% overall sealing effectiveness. 

There was no effect of the bluff bodies in the case of 1.5% purge flowrate. 

Figure 27. Pitch-average sealing effectiveness at the rim seal exit for the 

baseline rim seal with and without bluff bodies

NASA/CR—2015-218854 21
Figure 28. Nondimensional ingress levels throughout the baseline seal with bluff bodies present

Hot gas ingress levels were calculated using carbon dioxide (CO2) concentration measurements 

throughout the baseline seal, shown in Figure 28. The purge flow was seeded with 1% CO2 

(based on molar concentration) and the seal flow was sampled through the ports labeled 2 

through 6 in Figure 17. A nondimensional gas concentration was calculated as defined below: 

Ʉୡ ൌ ɖୡ െ ɖ

ɖୡ െ ɖஶ

An Ʉୡ value of 0 represents a high level of ingress where a value of 1signifies no ingress. 

Locations deep inside the seal (taps 2 and 3) showed consistently low levels of ingress. Locations 

near the trench (taps 4 through 6) showed the highest levels of ingress. As purge flow rate 

increased, the ingress near the trench decreased until all of the taps sampled yielded non 

dimensional gas concentration values higher than 0.7. The sensitivity of the region near the seal 

exit to increases in purge flow rate confirms that the instrumentation in the cascade is optimally 

placed to measure the impact of the main gas path pressure field on seal performance. 

Effects of Bluff Body Clocking Position on Sealing 

Effectiveness and Adiabatic Endwall Effectiveness 

An important question addressed by this work is that of what impact does the position of the 

downstream bluff bodies relative to the vanes have on rim seal performance. In an effort to 

answer this question, the location of the bluff bodies was shifted from their nominal clocking to 

two different alternate locations. The first was to the left by one third of the bluff body pitch. 

This location is referred to as the “+33% Clocking” position. The second location is one third of 

a bluff body pitch to the right of the nominal clocking location. This is referred to as the “-33% 

Clocking” location. These locations are shown in Figure 29. 

NASA/CR—2015-218854 22
Figure 29. Bluff body positions for all three clocking locations considered

The different locations of the bluff bodies change how their pressure field interacts with that of 

the upstream vane row. This interaction can be seen in the mid-span pressure distributions 

around each bluff body, shown in Figure 30. The +33% Clocking position shows the most 

uniformity between the three bluff bodies. This variation in pressure is related to how the wakes 

of the upstream vanes impinge on each bluff body. This reduced level of variation in the “+33% 

Clocking” position indicates that at that position the bluff bodies are removed from the vanes 

wakes and encounter uniform freestream flow. 

 (a) (b) (c) 

Figure 30. Pressure distribution from bluff body mid-span for the three clocking positions: 

(a) -33% pitch offset, (b) nominal position, (c) +33% pitch offset

Figure 31: Effects of bluff body clocking on average sealing effectiveness for the baseline rim seal

NASA/CR—2015-218854 23
Sealing effectiveness and adiabatic endwall effectiveness measurements were made to determine 

the effect of bluff body clocking on the baseline rim seal performance. Figure 31 shows the 

overall sealing effectiveness for the three clocking positions and at all three purge rates. The 

greatest variation occurs at the lowest purge rate. This is due to the rim seal becoming more 

insensitive to external influences with increasing purge rates. Even the greatest variation for the 

baseline rim seal, however, is still within or very near to the uncertainty of the measurement. The 

clocking position of the bluff bodies does not have a significant impact on the baseline rim seal. 

Bluff body clocking effects on two different rim seals were studied. These two rim seals are 

diagramed in Figure 32. The first is the baseline rim seal. The second incorporated two endwall 

extensions. One of which was attached to the trailing edge of the vane endwall and the other to 

the leading edge of the blade endwall. This modification extends each endwall by 2.5% of the 

vane axial chord length. This rim seal modification is called the short combined extension or 

(SCE) and doubles the axial overlap of the rim seal exit. 

Figure 32: Schematic showing the baseline rim seal and the geometry 

changes included in rim seal modification SCE

Bluff body clocking position had a greater impact on the SCE than on the baseline rim seal. The 

overall sealing effectiveness for both rim seals is shown in Figure 33. The solid lines correspond 

to the data from the baseline rim seal while the dashed lines show the data from the modified rim 

seal. The trends between the two rim seals are similar. The -33% Pitch Clocking position shows 

the lowest sealing effectiveness while the +33 Pitch Clocking exhibited the highest. These trends 

are clearly more exaggerated for the modified rim seal. The trends however are near the 

uncertainty of the measurement. The location of the bluff body had minimal impact on overall 

sealing effectiveness for both the baseline and the modified seal. 

NASA/CR—2015-218854 24
Figure 33: Average sealing effectiveness for the baseline rim seal and 

rim seal modification SCE with different bluff body clocking positions

The clocking effects on average sealing effectiveness at the lowest purge rate are better 

quantified in Figure 34. This figure shows the difference in average sealing effectiveness 

between the +33% Clocking and -33% Clocking positions from the nominal clocking position. 

The data from the nominal clocking position is denoted by the subscript “nom”. For the baseline 

rim seal the difference is within the uncertainty of the measurement, which is 0.02. The 

difference is greater for the data from rim seal modification SCE. These data show a difference 

of 0.05 overall sealing effectiveness between the %33% and -33% clocking positions. While this 

difference is greater than the uncertainty of the measurement it is less than 10% of the value 

from the nominal clocking location. 

The insensitivity of the baseline rim seal to bluff body position indicates that of the potential 

fields involved, that created by the vane row is dominant. This is supported by the increased 

variation in the data of the modified rim seal. By extending the vane endwall, the rim seal exit is 

moved away from the stronger regions of the vane potential field and closer to the potential field 

of the bluff bodies. This lessens the effect of the vane potential field and increases that of the 

bluff bodies on the rim seal. 

Figure 34: Effects of bluff body clocking on average sealing effectiveness 

for the baseline rim seal and rim seal modification SCE (0.75% purge rate)

NASA/CR—2015-218854 25
It is desired that through modification of the rim seal geometry that pitch-average sealing 

effectiveness may be increased. This augmentation in sealing effectiveness was evaluated at the 

rim seal exit for both rim seals. Bluff body clocking position was observed to have an impact on 

this augmentation. Figure 35 shows the augmentation in pitch-average sealing effectiveness due 

to rim seal modification SCE. The subscript 0 denotes data from the baseline rim seal. All three 

clocking positions and all three purges are shown. For the lower purge rates, 0.75% and 1% of 

the mainstream, an increase of 0.06 to 0.09 was observed respectively. The highest purge rate, 

1.5% of the mainstream, indicates no definite improvement. This is due to the fact that the 

baseline rim seal was already highly purged at this purge rate. The data from the purge rates 

where definite improvement in pitch-average sealing effectiveness was observed indicate that the 

+33% Pitch Clocking position tended to show slightly higher improvements. Bluff body clocking 

effects are smaller than the augmentation due to the rim seal modification considered. 

Figure 35: Increase in average sealing effectiveness due to 

rim seal modification SCE for various bluff body clockings and purge rates

The effect of bluff body clocking position on adiabatic endwall effectiveness was also 

considered. Adiabatic endwall effectiveness, defined below, was calculated from IR 

thermography measurements of the endwall temperature. Adiabatic endwall effectiveness values 

were averaged over the area outlined in red in Figure 15 to produce a metric to compare various 

seal geometries. Average adiabatic endwall effectiveness values for all rim seals, bluff body 

positions, and purge rates are shown in Figure 36. Uncertainty in adiabatic endwall effectiveness 

was calculated to be ±0.02. Bluff body clocking position shows little effect on the average 

endwall effectiveness. This indicates that the mixing in the trench is not greatly affected by bluff 

body clocking. 

Ʉ ൌ

ஶ െ ୟ୵

ஶ െ ୡ

NASA/CR—2015-218854 26
Figure 36: Effects of bluff body clocking on average endwall adiabatic 

effectiveness for the baseline rim seal and rim seal modification SCE

Trends in adiabatic endwall effectiveness augmentation due to rim seal modification were 

consistent with those from the sealing effectiveness data. This augmentation is shown in Figure 

37 for all clocking positions and purge rates. The “+33% Clocking” position consistently showed 

the highest increase. The augmentation at the “-33% Clocking” position was comparable to or 

less than that at the nominal clocking. 

Bluff body clocking position was varied by 33% of the bluff body pitch to either side of the 

nominal position. The effect of this variation on average sealing effectiveness at the rim seal exit 

and average adiabatic endwall effectiveness was studied. The variations in seal performance and 

adiabatic endwall effectiveness due to bluff body position were small. The insensitivity of the 

seal flow to bluff body position is mostly like due to the relatively small horse shoe vortex at the 

leading edge of the bluff body. The interaction between the seal flow and the horse shoe vortex 

at the leading edge of a blade has been shown in the literature to be sensitive to blade position. 

The bluff body was designed to match the pressure distortion of a rotating blade in a stationary 

reference frame. The leading edge radius of the bluff body is therefore much smaller than that of 

the blade, resulting in a smaller horse shoe vortex. The smaller vortex is not nearly as sensitive to 

bluff body position. 

NASA/CR—2015-218854 27
Figure 37: Effects of bluff body clocking on the increase in average 

endwall adiabatic effectiveness due to rim seal modification SCE

Parametric Study of the Rim Seal Geometry 

To better understand the trends associated with various rim seal geometry components, a 

parametric study of rim seal geometry was carried out. The effect of the axial overlap at the rim 

seal exit was the main focus of the study. Figure 38 shows the three rim seal modifications 

included in the study. The first modification consists of extending the blade endwall leading 

edge. The second represents an extension of the vane endwall trailing edge. The third 

modification is the combination of the first two. 

Figure 38: The relative location of the seal and the locations of the 

endwall extensions used in the parametric study on seal geometry.

NASA/CR—2015-218854 28
οX CΤ ܠ܉ οX۰۰ΤCܠ܉

Baseline 0.24 0.30

Seal Modification οVE/ο܁܀۳ οBE/ο܁܀۳

Baseline - -

Short Vane Endwall Ext. (SVE) 0.55 - 

Long Vane Endwall Ext. (LVE) 1.66 - 

Short Blade Endwall Ext. (SBE) - 0.55

Long Blade Endwall Ext. (LBE) - 1.66

Short Combined Ext. (SCE) 0.55 0.55

Long Combined Ext. (LCE) 1.66 1.66

Two different sizes of each rim seal modification were considered in the study and are listed in 

Table 4. The small combined extension increases the axial overlap at the rim seal exit by 110%. 

The larger combined extension increases the axial overlap by 330%. Thus the naming convention 

for the rim seal modifications consists of the short and long blade endwall extension (the S.B.E. 

and L.D.E. respectively), the short and long vane endwall extension (the S.V.E. and L.V.E. 

respectively) and the short and long combined extension (the S.C.E. and L.C.E. respectively). 

Extending the blade endwall serves to alter the axial momentum of the coolant in the seal. Due to 

the horizontal orientation of the rim seal exit, coolant with higher axial momentum will more 

effectively prevent the ingestion of main gas path fluid. The baseline rim seal and modification 

SCE have the minimum area of the seal in the vertical direction. Thus the coolant travels up and 

impinges on the underside of the vane endwall trailing edge in the vicinity of the rim seal exit. 

Impingement of coolant results in lower axial momentum as it mixes with gas from the trench at 

the rim seal exit. Increasing the blade endwall length directs the coolant into a hairpin turn 

preceded and followed by stretches of horizontal pathways before it reaches the rim seal exit. 

This increases the axial momentum of the coolant stream at the rim seal exit. Extending the vane 

endwall removes the source of ingested fluid from the rim seal. By extending the vane platform 

trailing edge the rim seal exit is distanced from the potential field of the vane which was earlier 

noticed to be the dominant in the ingestion process. Additionally, extending the vane endwall 

decreases the width of the trench through which ingress may occur. 

Extending the vane and blade endwall alters the flow patterns in the rim seal as well as the 

trench. The effects of these changes on pitch-average sealing effectiveness are illustrated in 

Figure 39 The data from the largest rim seal modifications are shown as they display the 

observed trends most clearly. The largest extension of the blade endwall, rim seal modification 

LBE, yields negligible sealing effectiveness gains over the baseline rim seal. This indicates that 

the influence of the trench flow field at the rim seal exit dominates the increase in axial 

momentum of the coolant. 

Significant gains were observed when the vane endwall was extended. This increase in sealing 

effectiveness arises from the effect of the vane potential field on the seal being mitigated and the 

decrease in axial width of the trench. Most interesting was the observation of a further increase 

in sealing effectiveness when the bluff body and vane endwalls were extended simultaneously 

(compare the LCE and LVE). This deviates from the negligible impact rim seal modification 

LBE had alone. The increase in axial coolant momentum caused by the blade endwall extension 

Table 4: Baseline Seal and Modification Parameters

may have a positive impact on sealing effectiveness at the rim seal exit when extending the vane 

endwall weakens the influence of vane row. 

NASA/CR—2015-218854 29
Figure 39: Average sealing effectiveness at the rim seal exit for rim seal modifications LBE, LVE, and LCE

The effect of each seal modification on sealing effectiveness is shown in Figure 40. The trends in 

the short extensions are similar to those of the long extensions previously discussed above. 

Those trends are most pronounced at the lowest purge rate (0.75% of the mainstream flow). As 

the purge rate decreases, seal design plays are larger role in preventing ingress. 

Figure 40: Improvement in sealing effectiveness relative to the baseline seal for all seal geometry modifications

The effect of rim seal modification on adiabatic endwall effectiveness was also considered. 

Figure 41 shows the effect of modification size on endwall effectiveness. The smaller 

modification showed only modest increase over the baseline rim seal. However by tripling the 

size of the endwall extensions, the average adiabatic effectiveness rose significantly. Figure 42 

confirms that the largest seal extensions yielded the largest increases in endwall effectiveness. 

By increasing the vane endwall extension from 2% vane axial chord to 6%, the average endwall 

effectiveness increases by a factor of 2 to 4 depending on purge rate and rim seal modification. 

Only at the highest purge rate did the short vane and combined endwall extensions (the SVE and 

SCE respectively) modestly increase cooling on the endwall. At the lower purge rates, the 

increase in endwall effectiveness was very near the uncertainty of the measurement (0.02 overall 

effectiveness). 

NASA/CR—2015-218854 30
Figure 41: Average endwall adiabatic effectiveness for rim seal modifications SCE and LCE

Figure 42 also shows that the cooling of the blade endwall was most affected by extending the 

vane endwall. The long vane endwall extension (the LVE) shows a considerable large increase 

than the long blade endwall extension (the LBE). Extending the vane endwall decreases the 

amount of ingress into the trench. As a result, the purge flow has a larger impact on the endwall 

temperatures. The long combined endwall extension (the LCE) yields even higher endwall 

effectiveness levels than the LVE. The higher momentum purge flow is reaching farther out into 

trench before mixing with the trench flow. 

Figure 42: Increases in average endwall adiabatic effectiveness for all rim seal modifications and all purge rates

A parametric study of the axial overlap at the exit of a radial overlap rim seal illustrates the 

relationships between the various endwall extensions and the performance of the rim seal. 

NASA/CR—2015-218854 31
Increasing the axial overlap by extending the blade endwall alone had minimal effect on the 

average sealing effectiveness at the rim seal exit and average adiabatic endwall effectiveness in 

the trench. Increasing the axial overlap by extending the vane endwall showed increased sealing 

and adiabatic effectiveness. Adiabatic effectiveness showed the greatest improvements when the 

largest vane endwall extension was used. The greatest gains in sealing and adiabatic 

effectiveness were obtained by simultaneously extending the vane and blade endwalls 

simultaneously. 

Flow field measurements in the trench region of the rim seal 

Flow field measurements in the trench region of the seal were made to illustrate the mechanisms 

behind the trends in seal performance previously reported. A key goal is to understand the 

mechanism through which hot gas from the main gas path enters the trench and then ingresses 

into the seal. Measurements of the three component velocity field were made using Laser 

Doppler Velocimetry. Measurements were made at two pitchwise locations in the trench, 

identified in Figure 43a. The first plane was located at the mid-pitch locations between the 

trailing edges of Vanes 3 and 4 and will be hereafter referred to as the mid-pitch plane. The midpitch plane coincides with an area of low endwall pressure as demonstrated in Figure 24. The 

second plane was placed in a region of high endwall pressure behind the trailing edge of Vane 4. 

The second plane was labelled as the vane trailing edge plane. Each plane covered axial and 

radial dimensions shown in Figure 43b. The locations of the two measurement planes facilitate 

the study of the changes in the trench flow field due to the passing of the vane wake. The wake 

of vane 3 had yet to cross the trench at the mid-pitch plane. The vane trailing edge plane 

corresponds to the location where the wake from vane 3 has nearly completely traversed the 

trench. 

(a) (b) 

Figure 43: Description of the planes over which flow field data were 

taken with the (a) pitchwise locations and (b) the radial and axial dimensions.

The flow fields at the mid-pitch and vane trailing edge planes are shown in Figure 44a and 

Figure 44b respectively. Vectors of in-plane velocity are shown with contours of pitchwise 

velocity normalized by the cascade inlet velocity. The quality of the data is supported by the 

good agreement between the normalized velocity in the main gas path (approximately four) and 

NASA/CR—2015-218854 32
the area ratio of the vanes in the cascade. The flow in the trench is predominately in the 

pitchwise direction due to the large amount of turning of the vanes. The purge flow was 

introduced into the seal with little pitchwise momentum. As a result, the regions of lowest 

pitchwise velocity in Figure 44 are close to the seal exit. Therefore fluid in areas of high 

pitchwise momentum must come from the main gas path. 

The flow at the mid-pitch plane, Figure 44a shows flow traveling along the blade endwall 

towards the rim seal exit. Partial ingress of the trench flow is observed at the seal exit. The fluid 

that does not enter the seal exit is turned up the vane endwall trailing edge by the purge flow. A 

shear layer is observed at the interface between the main gas path and the trench flow but does 

not extend more than three percent of the vane exit span into the main gas path. Figure 44b 

shows a distinctly different flow field at the vane trailing edge plane. Notably, the shear layer 

observed at the mid-pitch plane is replaced by a large region of low momentum fluid in front of 

the bluff body leading edge. The contours shown in Figure 45 indicate that removing the bluff 

body significantly decreases the low momentum fluid in the main gas path. This region of low 

momentum fluid extends more than ten percent of the vane exit span. The presence of low 

momentum fluid in the main gas path has a negative impact on aerodynamic efficiency. Fluid 

from the main gas path is also drawn into the trench at the trailing edge of the vane endwall due 

to the relatively high pressure at the vane trailing edge. 

 (a) (b) 

Figure 44: Contours of normalized pitchwise velocity with in-plane vectors for the 

baseline seal at (a) the mid-pitch plane and (b) the vane platform trailing edge plane.

NASA/CR—2015-218854 33
 (a) (b) 

Figure 45: Contours of normalized pitchwise velocity with in-plane vectors for the vane trailing edge 

plane for the baseline seal (a) with the bluff body and (b) without the bluff body. 

Turbulence plays an important transport role in the trench. Levels of turbulent kinetic energy 

(TKE) are shown in Figure 46. High levels of TKE are observed where the trench flow interfaces 

with the main gas path. Turbulence is produced in regions of gradients in mean velocity resulting 

in the highest levels of TKE. The unsteadiness of the enwall flows from the upstream vane 

further increases TKE in this region. Turbulence is transported throughout the entire trench 

region. Profiles of TKE in the trench shown in Figure 47 indicate that turbulence in the trench is 

as much as an order of magnitude larger than levels in the main gas path. Turbulence intensity 

levels were between 25 and 75 percent of the local velocity. The higher levels occurred near the 

seal exit due to the lower velocity associated with the purge flow. The significant turbulent 

mixing accounts for the low gradients of endwall adiabatic effectiveness shown in Figure 26. 

Turbulence mixes hot gas from the main gas path with the purge flow. The resulting mixture is 

then transported through the seal exit as indicated by the vectors in Figure 44. 

 (a) (b) 

Figure 46: Contours of normalized turbulent kinetic energy for the baseline seal at 

(a) the mid-pitch plane and (b) the vane platform trailing edge plane.

NASA/CR—2015-218854 34
Figure 47: Profiles of turbulent kinetic energy from the baseline seal 

A parametric study of the axial overlap at the seal exit was used to identify the effect of seal 

geometry modification on the transport mechanisms of hot gas ingress. Flow field measurements 

were made for the baseline seal as well as the LCE and LVE modifications. As the trends 

between the LCE and the LVE are similar, data from the LCE are compared with those from the 

baseline seal. The mean flow fields for the trailing edge plane of the baseline and LCE are shown 

in Figure 48. Extending the vane endwall reduced the tangential momentum of the fluid in the 

trench. The diminished effect of the vane pressure field and the decreased size of the trench 

width prevented hot gas from being drawn into the trench as observed with the baseline seal. 

Low momentum fluid is still pulled out into the mainstream. 

NASA/CR—2015-218854 35
 (a) (b) 

Figure 48: Contours of normalized pitchwise velocity from the vane trailing edge 

plane of the (a) baseline seal and (b) the L.C.E. seal. 

Extending the vane endwall also altered the turbulence in the trench flow field. Contours of TKE 

are shown in Figure 49 for the baseline seal as well as the LCE. The location of peak turbulence 

in for the LCE is shifted to the back of the trench. Extending the vane endwall altered the 

trajectory of the vane endwall flows. The unsteadiness of these flows may account for the high 

TKE levels near the bluff body leading edge. The turbulence near the seal exit in the LCE was 

diminished relative to the baseline seal. The profiles of TKE shown in Figure 50 indicate that the 

turbulence levels are somewhat comparable between the two seals. The gain in seal performance 

observed in the LCE and the LVE is due to the decrease in fluid being drawn into the trench as 

the vane wake crosses the trench and not due to diminished turbulent mixing. 

 

 (a) (b) 

Figure 49: Contours of normalized turbulent kinetic energy from the vane trailing 

edge plane of the (a) baseline seal and (b) the L.C.E. seal.

NASA/CR—2015-218854 36
Figure 50: Profiles radial velocity from the vane trailing edge plane of the baseline and L.C.E. seal. 

A cascade was used to study the effect of the main gas path pressure field on rim seal flows. The 

pressure field of both an upstream vane and a downstream blade were included in the study. The 

unsteady pressure field of a downstream blade was modelled using a bluff body. Bluff body 

clocking position had little effect on sealing effectiveness and adiabatic endwall effectiveness. 

The bluff bodies captured the pressure field of the blade but exhibited much smaller endwall 

flows than would exist in front of a rotating blade. The blade endwall flows have been shown to 

vary in size and interact with rim seal flows based on blade clocking position. The diminished 

endwall flows in front of the bluff bodies were too small to impact the flow in the trench region 

of the seal. 

A parametric study of seal geometry studied the impact of the axial overlap at the seal exit on 

seal performance. Extensions on the vane endwall, the blade endwall, and both endwalls 

simultaneously were used in the study. Extending the vane endwall increased sealing 

effectiveness as well as endwall cooling more than the extension of the blade endwall. The main 

mechanisms behind these improvements were the diminished impact of the vane pressure field 

and the decrease in trench width. A slight trend was observed that extending both endwalls 

simultaneously performed better than extending either alone. 

Flow field measurements in the trench region of the seal were made using a 3-component LDV 

system. Keys features of the trench flow field include recirculation pockets downstream of the 

vane trailing edge. A significant shear layer was observed varying in height between 5% and 

10% vane exit span. Low momentum fluid was observed to build up in front of the bluff body. 

Significant levels of turbulence kinetic energy were observed in the trench of all of the seals 

studied. The turbulence in the trench was enhanced by the unsteadiness of the vane endwall 

flows passing overhead of the trench. Extending the vane endwall altered the path of the vane 

endwall flows resulting in the peak TKE of the LCE to be located at the back of the trench. 

NASA/CR—2015-218854 37
3 AFTRF Rotating Rig 

Detailed descriptions of the overall facility, instrumentation configuration, and testing methods 

can be found in the Appendix section of this report. A brief summary will be provided here. The 

AFTRF facility had a pre-existing vane, blade, and rim cavity installed that were replaced with 

P&W hardware for the purposes of this NRA contract. Existing rig instrumentation ports, flow 

supplies, and measurement windows were re-used for the current work. Further hardware 

modifications to the AFTRF rig consisted of the alteration and installation of a purge chamber 

that was used to introduce purge fluid into the rim cavity. Modifications were made to a chamber 

that had been built for a previous DOE sponsored research program. The current leakage flow 

system is driven from a high pressure (300 psi) air reservoir. Up to 1.5% leakage flow mass flow 

rate is available in the current arrangement. The mass flow rate is monitored by a calibrated 

ASME orifice. Figure 51shows a photo of the AFTRF after the new P&W hardware was 

installed. 

Figure 51: AFTRF Facility With P&W Hardware For NRA Testing 

NASA/CR—2015-218854 38
Figure 52: AFTRF Purge Chamber (green and orange components) 

Inserted Below the Rim Cavity Region (purple and grey components) 

A cross section of the purge flow chamber and rim-cavity seal exit are shown in Figure 52. A 

non-rotating, flexible rubber seal was installed on the purge chamber to prevent leakage fluid 

from exiting the cavity before entering the main gas path flow. This rubber seal is represented in 

Figure 52 as the yellow component. A photograph of the installed hardware is on the left in 

Figure 53. The outer radius of the seal rubs against the rotating side of the rim cavity (gray 

component in cross section). 

 

SEAL 

Purge 

Chamber 

Installed on 

Vanes 

Figure 53: AFTRF Purge Chamber with the Seal Attached 

NASA/CR—2015-218854 39
AFTRF Instrumentation 

The AFTRF started to generate baseline data in November 2012 after hardware delivery to the 

laboratory was completed. Below is a summary of the installed measurement 

capability/systems: 

x NGV inlet flow traverse system 

x AFTRF Facility measurement control system and its computer/data interface 

x 150 channel slip-ring based data transmission system 

x NGV exit computerized flow traverse system 

x Rotor phase-locked measurement system 

x A Kiel probe based local total pressure measurement system 

x A five hole probe calibration jet for AFTRF class measurements 

x A time accurate total pressure measurement probe using a 150 Khz Endevco sensor 

(including its optical encoder based data collection/processing system) 

x Time accurate static pressure measurements using a 150 Khz Endevco sensors in the rimcavity 

A detailed view of the inlet measurement pitot probe is shown in Figure 54. The circumferential 

locations of the inlet probes are shown on the left and a picture of an inlet probe is on the right. 

During testing, the pneumatic average of the three probes was used to determine the inlet total 

pressure to the rig. 

NASA/CR—2015-218854 40
Figure 54: AFTRF Inlet Measurement Section (mid-span and full span traverses at one chord upstream of the NGV) 

Three of the inlet vanes were each instrumented with surface static pressure taps at three 

spanwise locations: 10%, 50%, and 90%. Surface static pressure measurements were also 

available on two of the rotor airfoils. These radial positions were 10%, 25%, 50%, 75%, and 

90%. 

Figure 55 shows a close-up of the purge exit region. Shown at this location is the insertion of a 

sub-miniature Kiel probe for the detailed measurement of vane exit quantities. The 

circumferential extent of these exit traverse planes was one vane pitch. Radial extents were 

between 10% to 90% span. A five-hole probe was also used to obtain flow exit angles and 

velocities.

300° 60°

180°

NASA/CR—2015-218854 41
Figure 55: Exit area of the purge chamber 

The AFTRF rotor also houses a rotating probe traversing mechanism, in addition to a 32 channel 

electronic pressure scanner. Connections to a 150 channel slip-ring system were re-designed and 

re-assembled to achieve a higher system reliability and ease of dis-assembly. This 

instrumentation is shown in Figure 56. 

 Figure 56: AFTRF Slip-Ring Instrumentation. 

A rotating probe traversing mechanism that can move an aero/thermal probe in both radial and 

circumferential directions was operated in AFTRF in Spring 2013. This traverser system allows 

a “high measurement grid resolution” for rotating frame measurements. A schematic of this 

measurement system is shown in Figure 57. Although much effort was given to the installation 

of this measurement system, useful data was never recorded and the program made the decision 

to forgo this measurement so that more geometries could be tested. 

BLADE 

NGV 

MAINSTREAM 

FLOW 

NASA/CR—2015-218854 42
Figure 57: AFTRF Rotating Measurement Schematic. 

A stationary-frame radial-circumferential traverse was used as a replacement measurement 

technique in place of the rotor-relative results that had been included in initial testing plans. Total 

pressure measurements were acquired with a Kiel probe and a high frequency response pressure 

probe manufactured by Endevco. When the latter probe results were post processed with a phaselock averaging method, the total pressure distribution for each individual rotor passage was 

visible. Distinct fluid structures were observed (e.g. tip vortex, etc.) which enabled a detailed 

analysis of how fluid structures changed as a function of hardware design variation and rim 

cavity purge flow rate. This probe proved to be a reasonable replacement for what would have 

been obtained with rotor-relative measurement techniques. 

Measurement uncertainty was estimated for each of the tranducers described above. Estimates of 

these uncertainties are listed below. Validyne transducers are used for measuring probes in the 

stationary frame of reference. Thermocouples are used to measure inlet and exit total 

temperature. Atmospheric pressure is measured using a gage on the wall. The Endevco Probe 

measures unsteady pressures behind the rotor from the stationary frame. The ZOC22b are used 

to measure blade and vane loading. The transducer with a one psi operating range is used to 

measure vane loading, and the more robust 2.5 psi unit measures blade loading from the rotating 

frame. Probe measurement uncertainties are tabulated in Table 5. 

NASA/CR—2015-218854 43
Validyne DP15-30 േͷܲܽ

K-Type Thermocouple േͲǤʹܭ

Atmospheric Pressure േʹͷܲܽ

Endevco േͷܲܽ

ZOC22b-1psi േͷܲܽ

ZOC22b-2.5psi േͳͷܲܽ

Moving downstream, the first measurement taken is the inlet velocity. This measurement is used 

calculate flow coefficient. Inlet velocity is a function of total pressure, static pressure, 

atmospheric pressure, and temperature. The equation for it is as follows: 

ܸ݅݊ ൌ ඨʹሺܲܶ െ ܲܵሻܴܶ

݉ݐܲܽ

The difference between total and static pressure is measured using one Validyne, meaning that 

the difference can be treated as one variable. To calculate uncertainty the following equation 

must be solved: 

߲ܸ݅݊ ൌ ඨቆ߲ሺܲܶ െ ܲܵሻ

߲ܸ݅݊

߲ሺܲܶ െ ܲܵሻ

ቇ

ʹ

൅ ቆ߲߲ܸܶ݅݊

߲ܶ ቇ

ʹ

݉ݐ߲ܲܽቆ൅ 

߲ܸ݅݊

ቇ݉ݐ߲ܲܽ

ʹ

߲ሺ்ܲ െ ܲௌሻ ൌ േͷ

߲ܸ݅݊

߲ሺܲܶ െ ܲܵሻ ൌ ܴܶ

ܲܽݐ݉ඨʹܴܶሺܲܶ െ ܲܵሻ

݉ݐܲܽ

߲ܶ ൌ േͲǤʹ

߲ܸ݅݊

߲ܶ ൌ ܴሺܲܶ െ ܲܵሻ

ܲܽݐ݉ඨʹܴܶሺܲܶ െ ܲܵሻ

݉ݐܲܽ

߲ܲ௔௧௠ ൌ േʹͷ

߲ܸ݅݊

݉ݐ߲ܲܽ

ൌ െܴܶሺܲܶ െ ܲܵሻ

݉ݐܲܽ

ʹ ඨʹܴܶሺܲܶ െ ܲܵሻ

݉ݐܲܽ

Inserting values average values into these equation results in a calculated uncertainty. The 

average amount of error for the inlet velocity is േͲǤ͵ ݉Τݏ. 

The next calculation is the vane loading. Since the vanes use the same non-dimensionalization, 

only one set of calculation will be shown. The difference between the two is only in magnitude 

of the measurements. Shown below is the equation for the coefficient of static pressure on the 

vanes and blades. This equation can be simplified by using Bernoulli’s principle for the velocity. 

Table 5. Measurement uncertainty for each probe used in this study. 

Measurement Uncertainty

NASA/CR—2015-218854 44
݉ݐܲܽ െ ܲܵ ൌ ݅ܿݐܽݐݏǡܲܥ

ͳ

ʹ

ൗ ߩܸ ʹൌ ܲܵ െ ܲܽݐ݉

ͳ

ʹ

ߩ ൗ

ʹሺܲܶ െ ܲܵሻ

ߩ

ൌ ܲܵ െ ܲܽݐ݉

ܲܶ െ ܲܵ

Both the denominator and the numerator are measured with their own respective Validyne. 

Meaning the equation is only a function of two measurements. To find the uncertainty the 

following equation is used: 

߲ܥ௉ǡ௦௧௔௧௜௖ ൌ ඨ൬߲ሺܲௌ െ ܲ௔௧௠ሻ ߲ܥ௉ǡ௦௧௔௧௜௖

߲ሺܲௌ െ ܲ௔௧௠ሻ

൰

ଶ

൅ ൬߲ሺ்ܲ െ ܲௌሻ ߲ܥ௉ǡ௦௧௔௧௜௖

߲ሺ்ܲ െ ܲௌሻ

൰

ଶ

߲ሺܲௌ െ ܲ௔௧௠ሻ ൌ ߲ሺ்ܲ െ ܲௌሻ ൌ േͷ

݅ܿݐܽݐݏǡܲܥ߲

߲ሺܲܵ െ ܲܽݐ݉ሻ ൌ ͳ߲ሺܲܶ െ ܲܵሻ

݅ܿݐܽݐݏǡܲܥ߲

߲ሺܲܶ െ ܲܵሻ ൌ െሺܲܵ െ ܲܽݐ݉ሻ

ሺܲܶ െ ܲܵሻʹ

Solving the equation for the vanes gives an uncertainty of േͲǤͷ, while solving for the blades 

results in an uncertainty of േͲǤ ͅ. 

Coefficient of total pressure used for both the Endevco and Kiel type probes behind the rotor, 

and refered to as ܥܲǡݐ

כ within this text, is given by the following equation. 

ݐǡܲܥ

כ ൌ ܲܶǡ݈݋݈ܿܽ െ ܲܶǡ݅݊

ܲܵǡ݁ݔ݅ݐ െ ܲܶǡ݅݊

ǡ௧௉ܥ߲

כ ൌ ඨቆ߲൫்ܲǡ௟௢௖௔௟ െ ்ܲǡ௜௡൯ ߲ܥ௉ǡ௧

כ

߲൫்ܲǡ௟௢௖௔௟ െ ்ܲǡ௜௡൯

ቇ

ଶ

൅ ቆ߲൫ܲௌǡ௘௫௜௧ െ ்ܲǡ௜௡൯ ߲ܥ௉ǡ௧

כ

߲൫ܲௌǡ௘௫௜௧ െ ்ܲǡ௜௡൯

ቇ

ଶ

߲൫்ܲǡ௟௢௖௔௟ െ ்ܲǡ௜௡൯ ൌ ߲൫ܲௌǡ௘௫௜௧ െ ்ܲǡ௜௡൯ ൌ േͷ

ݐǡܲܥ߲

כ

߲൫ܲܶǡ݈݋݈ܿܽ െ ܲܶǡ݅݊൯

ൌ ͳ

ܲܵǡ݁ݔ݅ݐ െ ܲܶǡ݅݊

ݐǡܲܥ߲

כ

߲൫ܲܵǡ݁ݔ݅ݐ െ ܲܶǡ݅݊൯

ൌ െ ܲܶǡ݈݋݈ܿܽ െ ܲܶǡ݅݊

൫ܲܵǡ݁ݔ݅ݐ െ ܲܶǡ݅݊൯

ʹ

To find the uncertainty, the worst case scenario is chosen by trying the min and max numbers. 

But since the numbers are so close together (about 100 pascal difference between the Endevco 

and Kiel) the resulting uncertainty comes out to be the same. For both the resulting uncertainty 

ݐǡܲܥ in

כ is േͲǤͲͲʹ

The Five-Hole Probe’s uncertainty is calculated in much the same way. The derived total and 

static pressure measurements are a function of the five pressure measurements and a calibration 

NASA/CR—2015-218854 45
factor. The results of the calculations are presented in Table 6 below, along with the rest of the 

results of the calculated uncertainties. 

Table 6. Derived uncertainty for each calculated variable shown in this study. 

Measurement Uncertainty Percent

࢔࢏ࢂ േͲǤ͵ ݉Τݏ േͳǤ ͅ ͅΨ

ࡼ࡯ǡࢉ࢏࢚ࢇ࢚࢙ Vane േͲǤͷ േ͵ǤʹͶΨ

ࡼ࡯ǡࢉ࢏࢚ࢇ࢚࢙ Blade േͲǤ ͅ േʹǤͻͻΨ

࢚ǡࡼ࡯

כ Kiel േͲǤͲͲʹ േͲǤʹͲΨ

࢚ǡࡼ࡯

כ Endevco േͲǤͲͲʹ േͲǤʹͲΨ

ࢀࡼ FHP േͳͲܲܽ േͲǤͷͲΨ

ࡿࡼ FHP േͳͲܲܽ േͲǤͷͲΨ

ࢂ FHP േͲǤ͹ͷ ݉Τݏ േͳǤͲ͹Ψ

 

NASA/CR—2015-218854 46
AFTRF RESULTS 

Vane loadings 

Static pressure taps are at three different spanwise locations on three of the inlet vanes and five 

locations on two of the blades (both baseline and TVC). They are at a constant radius 

corresponding to 10%, 50%, and 90% span on the vane and 10%, 25%, 50%, 75%, and 90% 

span on the blades. The equation below shows the definition for static pressure coefficient for the 

following plots. 

ܥ௉ǡ௦௧௔௧௜௖ ൌ ܲ௦ െ ܲ௔௧௠

ͳ

௡௜ܸ௡௜ߩ ൗʹ

ଶ

Each pressure plot also contains a numerical prediction provided by P&W. The numerical 

simulations were conducted with an in-house RANS solver utilizing a laminar to turbulent 

transition model to capture boundary layer transition behavior. 

Figure 58 shows a plot of Cp for the vane at the 10% span location for one of the inlet vanes. The 

results between the measured and the CFD prediction match fairly well in these cases. The 

pressure side (upper curve) matches more closely than the suction side (lower curve) of the 

vanes. 

Figure 58: Static Pressure Vane Loading at 50% Span

Blade Loadings 

Static pressure loading on the blades is a more challenging measurement than the vane loading. 

Blade loading must contend with centrifugal pumping in the tubing, pressure drop within the 

drum, and centrifugal loading on the pressure transducers. Corrections for each of these factors 

are made to all of the measurements taken in the rotating frame. 

NASA/CR—2015-218854 47
Figure 59 shows a plot of Cp from the baseline blade at 10% span. Agreement with the 

numerical predictions for the suction and pressure surfaces is generally good. At each static tap 

location there are two symbols representing the two instrumented blades installed on the rotor. 

At this spanwise position, the two blades show good agreement with each other. 

Figure 59: Baseline Rotor Blade Loading at 10% Span

Intraspace Measurements 

Vane exit surveys were conducted with a five hole probe and Kiel probe. Positive angle 

conventions are needed to understand what the FHP is measuring. Figure 60 show that a positive 

axial direction is downstream, a positive theta direction is counterclockwise, and a positive radial 

direction is toward the center of rotation. Figure 61 shows the positive angles and the planes they 

are working in in a 3⁄4 view. Positive alpha would be coming from below the FHP as shown in 

the blue meridional plane, and positive beta will be coming from the left of the probe shown in 

the green radial/theta plane. The rotor is represented by the purple oval. 

NASA/CR—2015-218854 48
Figure 60: Five-Hole Probe Positive Conventions, Downstream of Rotor Looking Toward Inlet

Figure 61: Five-Hole Probe Positive Angle Conventions

Vθ

VRadial

VAxial

(Out of page)

Positive Alpha

Positive Beta

Probe in Rig

VAxial positive downstream

Vθ positive counter clockwise

VRadial positive center of rotation

α positive when flow is moving in positive radial direction

β positive when flow is moving in positive θ direction

NASA/CR—2015-218854 49
Figure 62 shows total pressure coefficient measured from the Kiel at this station. The wake of 

the vane is clearly visible in the left side of the contours. The passage vortex impact on total 

pressure loss is visible below 20% span, centered at +1 deg. On the OD endwall, secondary flow 

impact is visible in the circumferential non-uniformity in boundary layer development. 

Figure 62: Five-Hole Probe Comparison without Rim-Seal Chamber Installed

 

NASA/CR—2015-218854 50
Five-Hole Probe Measurements at Vane Exit 

Figure 63: Coefficient of Total Pressure in the Intraspace 868 Point Mesh Measured by a Five-Hole Probe

Figure 63 shows a contour plot of total pressure coefficient as measured by the five hole probe. 

The flow field corroborates what was measured by the Kiel probe (Figure 62) in that the airfoil 

wake and endwall structures are clearly identified and in the same positions. The FHP does 

measure a slightly wider airfoil wake in the region between 20% and 80% span. 

 

Above 90% span is a large region of loss. It is noted that in this region is the window slot for 

probe access. It is the opinion of the researcher that some of the loss here is due to improper 

sealing. Improper sealing could induce flows on the probe that are out of acceptable ranges 

(±30°) on the FHP. A second reason could be due to proximity to the outer casing. The final 

reason could be due to the fact that the probe is fully or partially recessing into the cavity for 

these measurements. The large pressure drop in this region could be caused by the probe no 

longer being completely exposed to the main gas path. 

Figure 64 shows the static pressure coefficients at the vane exit plane. Lower static pressure 

coefficient corresponds to a lower static pressure. The radial distribution characteristic of a 

swirling annular flow is clearly visible. Circumferential variations in static pressure are a result 

of the vane potential field extending downstream of the vane trailing edge. 

NASA/CR—2015-218854 51
Figure 64: Coefficient of Static Pressure in the Intraspace 868 Point Mesh Measured by a Five-Hole Probe

 

NASA/CR—2015-218854 52
Blade Exit Measurements 

The baseline blade was used to demonstrate the new rig hardware. Measurements downstream of 

the blade were taken with a Kiel probe and with an Endevco probe. The measured total pressure 

coefficient, Cp,total* is defined as: 

ܥ௉ǡ௧௢௧௔௟

כ ൌ ܲ௧ǡ௟௢௖௔௟ െ ܲ௧ǡ௜௡

ܲ௦ǡ௘௫௜௧ െ ܲ௧ǡ௜௡

Ǥ

Endevco measurements were taken at zero rim cavity purge flow rate. These results are shown in 

Figure 65. These contours of total pressure coefficient were, first, phase-lock averaged to obtain 

a full wheel measurement map of the full rotor, and then passage-averaged to eliminate 

circumferential measurement variations that were present in the Endevco results. This figure 

shows the final passage-averaged contours. 

Figure 65: Contours of average total pressure coefficient from the 

Endevco probe

 

NASA/CR—2015-218854 53
Purge Flow Impact on Rotor Exit: Kiel Measurements 

Purge flow was introduced into the rim-seal cavity at four different rates. The rates are carefully 

monitored using an ASME calibrated orifice and total pressure contour maps are taken for each 

case. The variation in purge flow rate for these tests is tabulated in Table 7. Purge rate 

percentages are relative to the main gas path flow rate. 

Table 7: Rim cavity purge flow rate test matrix

Blade Type Tip Style Purge Rate 

Baseline Smooth 0.00% 

0.25% 

0.50% 

1.00% 

These flows were measured with the stationary-frame Kiel probe. The Kiel was radially 

traversed downstream of the rotor. The resulting time averaged data is shown as radial profiles of 

total pressure coefficient in Figure 66. Increasing purge flow rate increases the secondary flow 

penetration height through the passage. 

NASA/CR—2015-218854 54
Figure 66: Baseline Blade Purge Varying Purge Rate

 

NASA/CR—2015-218854 55
4 Summary Of Contract Data Deliverables 

Task Description Status

1 Low Speed Cascade Facility Modification

New vane geometry Complete

Bluff body geometry which is clockable Complete

Provisions for upstream and downstream 

leakage Yes, only using downstream

2 Instrument LSC Hardware and Test Section

Optical access for LDV, PIV Yes

Static pressure taps Yes

5-hole pressure probe Yes

Surface TC's Yes

Thermocouple probes Yes

As required for measuring: -

Aero loss Yes, 5-hole probe

Steady Ps Yes, Ptaps

Unsteady Ps No, determined not needed

Rim cavity effectiveness IR thermography, TC rakes, CO2

3 AFTRF Modification

Modify existing baseline stage for receiving 

vanes, blades, and rim cavities as req'd Complete

Install rim cavity Complete

Design, build, assemble a new intra-stage 

traverse for detailed flow measurements Complete

Design, build, assemble a new rotor exit traverse 

that moves both radial and circumferential Complete

Modify rotating instrumentation package Complete

Improve and assemble a modified phase-locked 

and time accurate rotor exit total prssure 

mapping system

Complete

Design and build a new measurement system 

based on stage inlet and exit for measuring 

efficiency

Complete

4 Instrument AFTRF Hardware

Rotating frame measurements Yes

High response instrumentation Yes, Endevco

Hot wire Yes

Conventional probes Yes

Pressure taps in flowpath and rim cavity Yes

As required for measuring: -

Flowpath and rim cavity flow fields using high 

response radial/circumferential traversing 

probes

Yes

Steady and unsteady airfoil, platform, and rim 

cavity surface static pressures Yes

Rim cavity effectiveness (ingestion) No

NASA/CR—2015-218854 56
 

Task Description Status

5 LSC Testing

Experimental measurements on 12 or more 

configurations (including blade, vane, endwall, 

cavity, and clocking changes).

Complete

Measure aerodynamic losses Complete

Measure steady airfoil, platform, and rim cavity 

surface static pressures Complete

Measure rim cavity ingestion effectiveness Complete

6 Baseline AFTRF Test

Shake down and baseline geometry testing to be 

used for comparing optimized geometries tested 

later.

Complete

7 LSC Testing

Experimental measurments on 4 or more 

configurations (including blade, vane, endwall, 

cavity, and clocking changes).

(1) Baseline

(2) Baseline with bluff bodies

(3) Long-Combined Endwall

(4) Long Vane Endwall

Measure aerodynamic losses In-Progress

Flowfield measurements (PIV or LDV) Complete

8 AFTRF Testing

Experimental measurments on 3 or more 

configurations.

(1) Grooved baseline tip

(2) TVC Blade

(3) Grooved TVC tip

(4) New rim-seal geometry

Flowfield measurements using high response 

radial/circumferential traversing probes Endevco and Kiel

Steady and unsteady airfoil, platform, and rim 

cavity surface static pressures Complete

Rim cavity ingestion effectiveness No

NASA/CR—2015-218854 57
5 CONCLUSIONS 

LSC: 

- Trend-accurate low speed cascade. 

- Matches upstream and downstream pressure field of the rim-cavity exit. 

- A place to test concepts in a region of the engine where CFD solutions are still being 

developed. 

- Cascade results show that extending the vane trailing edge of the seal provides benefit 

and extending the blade leading edge of the seal does not provide benefit. Benefit is 

improved seal effectiveness (reduced ingestion and purge flow cooling on downstream 

platform). 

- Cascade results are insensitive to location of downstream bluff body 

AFTRF: 

- New low-speed rig geometry installed, matched CFD airfoil loadings 

- Increasing purge flow in low speed rig pushes ID loss from horseshoe vortex radially 

outward. 

 

NASA/CR—2015-218854 58
APPENDIX: TEST ARTICLE PHOTOS 

AFTRF 

AFTRF assembled vane ring. The cut-away in the case in the upper right corner is where the 

optical access port is located. 

AFTRF traverse and port cover. 

 

NASA/CR—2015-218854 59
AFTRF assembled blades and rotor. 

LSC 

LSC instrumented vane (static pressure tap leads) 

LSC vane endwall and upstream rim cavity, showing vane cut-away and thermocouple rakes 

installed in the rim cavity. 

 

NASA/CR—2015-218854 60
6 REFERENCES 

Gibson J., Thole K., Christophel J., Memory C., 2015 “Effects of the Main Gas Path Pressure 

Field on Rim Seal Flows in a Stationary Linear Cascade,” ASME Paper No. 2015-43517

NASA/CR—2015-218854 61