Mounir Ibrahim and Abhinav Gupta

Cleveland State University, Cleveland, Ohio

Vikram Shyam

Glenn Research Center, Cleveland, Ohio

Computational Investigation of Novel Tip

Leakage Mitigation Methods for High

Pressure Turbine Blades

NASA/TM—2014-218494

November 2014
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Mounir Ibrahim and Abhinav Gupta

Cleveland State University, Cleveland, Ohio

Vikram Shyam

Glenn Research Center, Cleveland, Ohio

Computational Investigation of Novel Tip

Leakage Mitigation Methods for High

Pressure Turbine Blades

NASA/TM—2014-218494

November 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

Part of this work was conducted by the lead author while he was a Summer Fellow at the NASA Glenn Research Center and with 

Dr. Shyam as the NASA colleage. NASA Glenn is greatly acknowledeged for such support. We greatly appreciate the computer 

time provided for us by the Ohio Super Computer (OSC). The OSC Computer Cluster has been made available as part of the 

Center’s mission to support Ohio Universities. 

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

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

either expressed or implied, by the National Aeronautics and 

Space Administration.

This work was sponsored by the Fundamental Aeronautics Program 

at the NASA Glenn Research Center.

Level of Review: This material has been technically reviewed by technical management. 
NASA/TM—2014-218494 1

Computational Investigation of Novel Tip Leakage Mitigation 

Methods for High Pressure Turbine Blades 

Mounir Ibrahim and Abhinav Gupta 

Cleveland State University 

Cleveland, Ohio 44115 

Vikram Shyam 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

This paper presents preliminary findings on a possible approach to reducing tip leakage losses. In this 

paper a computational study was conducted on the Energy Efficient Engine (EEE) High Pressure Turbine 

(HPT) rotor tip geometry using the commercial numerical solver ANSYS FLUENT. The flow solver was 

validated against aerodynamic data acquired in the NASA Transonic Turbine Blade Cascade facility. The 

scope of the ongoing study is to computationally investigate how the tip leakage and overall blade losses 

are affected by (1) injection from the tip near the pressure side, (2) injection from the tip surface at the 

camber line, and (3) injection from the tip surface into the tip separation bubble. The objective is to 

identify the locations on the tip surface at which to place appropriately configured blowing keeping in 

mind the film cooling application of tip blowing holes. The validation was conducted at Reynolds 

numbers of 85,000, 343,000, and 685,000 and at engine realistic flow conditions. The coolant injection 

simulations were conducted at a Reynolds number of 343,000 based on blade chord and inlet velocity and 

utilized the SST turbulence model in FLUENT. The key parameters examined are the number of jets, jet 

angle and jet location. A coolant to inlet pressure ratio of 1.0 was studied for angles of +30°, –30°, and 

90° to the local free stream on the tip. For the 3 hole configuration, 3 holes spaced 3 hole diameters apart 

with length to diameter ratio of 1.5 were used. A simulation including 11 holes along the entire mean 

camber line is also presented (30q toward suction side). In addition, the effect of a single hole is also 

compared to a flat tip with no injection. The results provide insight into tip flow control methods and can 

be used to guide further investigation into tip flow control. As noted in past research it is concluded that 

reducing leakage flow is not necessarily synonymous with reducing losses due to leakage. 

Nomenclature 

Cp   2 2 / PT  P UUe , pressure coefficient

Cx blade axial chord (132.59 mm) 

P pressure 

PS upstream static pressure 

PT upstream stagnation pressure 

Re Reynolds number based on inlet conditions and axial chord 

Ue nominal exit freestream velocity, based on inviscid solution 

Z spanwise coordinate 

X axial coordinate, axial distance from leading edge 

Y pitchwise coordinate 

U density 

\ (PT-PTe)/(PT-PS), total pressure loss coefficient 
NASA/TM—2014-218494 2 

Introduction 

NASA’s N+3 notional engines are expected to be characterized by high operating pressure ratios and 

turbine inlet total temperatures. The more compact highly loaded turbines will result in large pressure 

difference across the turbine rotor tips. The pressure distribution set up around the rotor blade results in a 

pressure gradient across the tip gap of the rotor. This is a significant concern due to the occurrence of 

aerodynamic loss due to tip leakage. The high pressure gas on the pressure side of the blade has a 

tendency to flow towards the suction side, which is at a lower pressure, across the tip. The path of the tip 

flow depends on several parameters such as blade rotational rate, blade geometry (tapering, twist, camber, 

tip gap height, and tip contouring) and flow inlet angle. These parameters can be reduced to the ratio of 

tip width to tip gap height, tip inlet Mach number and flow direction. No matter what the cause of the 

leakage flow, it results in a drop in efficiency and aerodynamic loss. The loss due to the interaction of the 

tip leakage vortex with the passage flow comprises approximately 20 to 30 percent of the total loss in a 

turbine stage. It is therefore important to reduce the leakage across the tip relative to the massflow 

through the passage. The width-to-height ratio of the tip gap changes as the chord increases from leading 

edge to trailing edge. This is accompanied by a change in the flow structure in the tip gap. If one is to 

attempt to control the flow through the tip gap, it is important to take into account the location and extent 

of the separation bubble, the local flow direction and the mass flow rate per unit chord as a function of the 

mass flow rate through the passage. A detailed literature review of the basic features of turbine tip flow 

and heat transfer has been conducted by Bunker (Ref. 1) and Glezer et al. (Ref. 2). Glezer et al. (Ref. 2) 

describe the primary flow features seen due to tip leakage: the pressure side separation bubble on the tip 

surface and the tip leakage vortex. The extent of the bubble and its reattachment are contingent on blade 

thickness, Mach number, Reynolds number and tip height (Refs. 2 and 3). For subsonic flows, the 

percentage of inlet flow that constitutes the tip leakage is shown to grow linearly with tip gap height. 

References 1 and 2 highlight the importance of Computational Fluid Dynamics (CFD) in tip flow and 

heat transfer prediction due to the difficulty of conducting experimental measurements in a rotating tip 

gap. Ameri et al. (Ref. 4) simulated the GE-EEE (Energy Efficient Engine) high pressure turbine stage 

and studied the tip flow and heat transfer for a smooth tip and for a recessed tip. The cavity height of the 

recess was varied. Two primary vortex structures were observed in the recess. It was found that the recess 

had negligible effect on loss. Another study by Ameri et al. (Ref. 5) investigated the effect of upstream 

casing recess on tip leakage and heat transfer and found that minimal tip heat transfer occurred for a 

recess height that is almost equal to tip clearance. They concluded that the recess has little effect on 

efficiency. Hofer et al. (Ref. 6) state that neglecting the relative casing motion does not significantly alter 

results citing previous studies like that of Krishnababu et al. (Ref. 7). The latter studied the effect of 

relative casing motion on tip heat transfer and tip leakage mass flow for two different tip gap heights. 

They concluded that the effect of relative casing motion is diminished for larger tip gap heights due to 

dominant inviscid effects. The flow physics in a supersonic tip gap such as one encountered in highly 

loaded turbine stages is described in Shyam et al. (Refs. 8 and 9) and Wheeler et al. (Ref. 10). The 

importance of choking the tip gap to reduce the relative massflow through the gap is emphasized by both 

authors. 

In the realm of tip leakage mitigation, there are numerous passive and active approaches. Passive 

approaches involve strategies like contouring of the tip and blowing into the tip and casing treatments. 

Active methods would involve real-time adjustments to the tip gap geometry or flow that can be varied 

based on mission conditions. This work deals with steady blowing into the tip gap and thus a brief 

literature review of relevant passive tip leakage control is presented. Rao et al. (Ref. 11) found through an 

experiment that injection into the tip of a recessed blade through 4 cooling holes reduced the losses 

associated with tip leakage. A coolant massflow of 0.6 percent of the passage massflow injected into the 

tip was found to be the global optimum as evaluated by averaged pressure loss coefficient in the wake. 

Couch et al. (Ref. 12) compared CFD to measurements of blowing from purge holes on the turbine blade 

tip. They found a strong dependence of cooling effectiveness on tip gap height. Studies of film cooling 

(such as Refs. 12 to 14) show the dependence of jet behavior on momentum ratio, injection angle, pitch 
NASA/TM—2014-218494 3 

and plenum conditions. Canelli et al. (Ref. 14) studied film cooling from holes in the recessed tip of a 

turbine blade. They simulated 3 holes near the leading edge region along the mean camber line and holes 

near the suction side further down the axial chord. Holes are generally placed along the mean camber line 

due to ease of manufacturing. It is clear that along the chord in the tip gap, the blowing ratio would 

change based on local conditions in the tip gap. There are several unanswered questions related to 

managing tip leakage using blowing. These relate to the optimal locations at which the coolant holes 

should be placed, the spacing of the holes, their inclination, the blowing ratio, frequency and the effect of 

loading (subsonic or supersonic flow) on the leakage control. Frequency is mentioned because it might be 

possible to modify the loss core with pulsed jets such as those used for separation control by Ibrahim 

et al. (Refs. 15 and 16). 

This paper examines the effect of the number of holes along the mean camber line and the effect of 

leading edge hole inclination on leakage induced losses. While the mean camber line might not be the 

optimal location for leakage mitigation, the intent is to determine the influence of the holes on the leakage 

through the tip and on the overall pressure loss through the passage. Cooling holes are usually placed in 

locations in the tip region where heat transfer is high and to purge dust that might accumulate in sandy 

environs (Ref. 13). They are also usually inclined at 90q to the tip surface and placed within squealer tips 

(Refs. 12 and 14). The scope of this paper is to answer the question of global impact of number of holes 

and of hole inclination on the overall loss across the turbine blade passage. Future investigations will look 

at heat transfer on the tip and the effect of blowing into the separation bubble as well as varied hole 

inclination along the tip as a function of chord. 

Numerical Methods 

The numerical simulations were conducted utilizing the finite-volume code ANSYS FLUENT 

(Ref. 18). For validation, cases were run at different Reynolds numbers. Table 1 shows the cascade 

parameters. The computational domain was based on Ibrahim et al. (Refs. 15 and 16). A correlation-based 

transition model proposed by Menter et al. (Ref. 18) that is available in ANSYS FLUENT V-12 as 

Transition-SST (4-equation) turbulence model was selected to model turbulence. Figure 1 shows the 

computational domain and the boundary conditions applied in this paper. The figure depicts a case with 

11 jets at the tip equally spaced between a distance from 0.13 to 0.9 Cx. The number of cells used is 

9.2 million with 50 cells packed in the tip gap region in the spanwise direction. The tip gap height is 

2 percent of the span. The grid size used was reached from a grid indepence study that was conducted on 

the same blade. Convergence was established when: (1) residuals reduced below a value of 10–5, (2) no 

change was observed in any field results, and (3) the mass imbalance between inlet and exit to the 

computational domain was less than 0.01 percent. 

Experimental Facility and Measurements 

The results of the numerical simulations that were carried out are compared to the data of McVetta 

et al. (Ref. 17). The data was collected in the NASA Transonic Turbine Blade Cascade facility at the 

NASA John H. Glenn Research Center. The data includes the midspan total pressure at the exit plane and 

the exit flow angle at seven inlet flow angles. Reynolds numbers of 85,000 to 683,000 at two isentropic 

exit Mach numbers of 0.74 and 0.34 are reported. The survey plane at midspan was 8.6 percent axial 

chord downstream of the blade trailing edge plane and covered three blade passages. Blade and endwall 

static pressure distributions were reported for each flow condition. Table 1 shows the relevant parameters 

used in the cascade experiment of Reference 17. 

Figure 2 shows the layout of the experiment in Reference 17. The turbulence intensities are reported 

to be between 0.25 and 0.4 percent.
NASA/TM—2014-218494 4 

Results and Discussion 

Code Validation and Turbulence Model/Geometry Selection, No Gap 

In the validation stage, the blade was assumed to have no tip gap and the results are compared to the 

data from Reference 17. A symmetry plane was used at mid-span after comparison to a full span 

simulation that showed no benefit to the additional computational domain. Shown below are static 

pressure distributions for two different turbulence models (RSM and SST) at Reynolds numbers, Re, of 

85,000, 343,000, and 685,000. Results at spanwise locations of Z = 50, 75, and 97.5 percent are 

presented. The pressure loss coefficient at a location X = 1.086 Cx from the airfoil leading edge and at 

Z = 50 percent is also compared to experiment. 

Re = 85,000 

Figures 3(a) to (c) show the static pressure coefficient plotted against dimensionless axial location, 

X/Cx for Z = 50, 75, and 97.5 percent respectively for Re = 85,000. The following observations were 

made: (1) The pressure distribution on the pressure surface of the airfoil are in close agreement with the 

experimental data regardless of the turbulence model used, or the Z location, (2) As for the suction 

surface the plots show some flow separation in the range of X/Cx from 0.65 to 0.78 after that the flow 

reattach, (3) The two turbulence models, RSM and SST, at Z = 97.5 percent showed flow separation as 

expected (there is no experimental data at that location to verify this although the wake shown in 

Figure 3(d) for the experiment shows a wide wake that is indicative of separation.) 

Figure 3(d) shows the total pressure distribution along X/Cx = 1.086 and Z = 50 percent plotted 

against y/pitch. The two turbulence models, RSM and SST, showed smaller wake width as compared to 

the experimental data perhaps indicating that the SST predicted separation is local and the flow 

reattaches. The RSM model does not show separation and therefore has a narrow wake. 

Re 343,000 

Figures 4(a) to (c) show the static pressure coefficient plotted against dimensionless axial location, 

X/Cx for Z = 50, 75, and 97.5 percent respectively, for Re = 343,000. The following observations were 

made: (1) The pressure distribution on the pressure surface and the suction surface of the airfoil are in 

close agreement with the experimental data regardless of the turbulence model used, or the Z location. 

Figure 4(d) shows the total pressure distribution along X/Cx = 1.086 and Z = 50 percent plotted 

against y/pitch. All cases exhibit similar wake size and peak value with SST turbulence model showing 

slightly better agreement with the experimental data. The absence of separation removes any 

discrepancies between the models. 

Re 685,000 

Figures 5(a) to (c) show the static pressure coefficient plotted against dimensionless axial location, 

X/Cx for Z = 50, 75, and 97.5 percent respectively, for Re = 685,000. Figure 5(d) shows the total pressure 

distribution along X/Cx = 1.086 and Z = 50 percent plotted against y/pitch. The results for Re = 685,000 

have similar observations shown earlier for Re = 343,000. 

Figure 6 shows a plot of the loss coefficient, area averaged defined as: 

in exit

in exit

total, static,

total, total,

P P

P P



 

for different Reynolds numbers comparing experimental data (Ref. 17) with CFD results. Figure 6 shows 

that both turbulence models underpredict loss at low Reynolds number (Re = 85,000) because of their 

failure to predict an open separation. At higher Reynolds numbers both models show improved 

predictions with the SST model doing a better job. 
NASA/TM—2014-218494 5 

Summary of Code Validation 

All cases examined in this section exhibit the following: 

1. The static pressure distribution on the pressure surface of the airfoil is in close agreement with the 

experimental data regardless of the turbulence model used, or the Z (span-wise) location. 

2. As for the suction surface, all cases exhibit similar results to the experimental data except at 

Re = 85,000, where the SST model shows a closed separation and the RSM model shows no 

separation. The experimental results suggest a wide wake that is not matched by the CFD. 

3. As for the total pressure results at a location 1.086 Cx from the leading edge, and Z = 50 percent, the 

two turbulence models, RSM and SST, showed good agreement with the experimental data for higher 

Re (343,000 and 685,000). The wake predictions for Re = 85,000 is not as good for the following 

reasons: (a) Transition is expected and was not simulated well with the turbulence models, (b) The 

simulation did not include the boundary layer in the inlet flow, (c) The span of the blade is smaller 

(approximately equal to the cord length), (d) literature review indicates the difficulty in predicting the 

wake for transitional flows. 

4. The overall conclusion is that there are good perditions of the experimental data especially at higher 

Re and with the SST model. 

In the rest of the paper analyses will be done for Re = 343,000 using the SST turbulence model. 

Tip Leakage With Different Jet Configuration 

Different configurations were examined to investigate the effect of various jet configurations on the 

tip leakage. 

The gap size is 2 percent Cx. All jets are located at the mean camber line, 5 mm in diameter and with 

jet length/diameter ratio of 1.5. Table 2 shows three different jet configurations examined: (1) One jet 

located at X/Cx = 0.23 measured from the leading edge, (2) Three jets equally spaced over a distance 

from X/Cx = 0.23 to 0.69 measured from the leading edge, (3) Eleven jets equally spaced over a distance 

from X/Cx = 0.13 to 0.9 measured from the leading edge. In addition different jet angles were examined: 

(a) 30q minus (i.e., aiming towards the suction surface), (b) 90q normal to the tip surface, (c) 30q plus 

(i.e., aiming towards the pressure surface). The jet inlet pressure was kept equal to the inlet total pressure 

(= 36.942 kPa) for all cases. Accordinlgy (as will be shown later) the jet blowing ratio (based on jet exit 

velocity/inlet velocity) will vary from 1.0 to 2.2 depending on the number of jets, jet location and jet 

orination. This is expected to similaute real life situation where all jets will be supplied from one plenum. 

Figures 7(a) to (c) show the static pressure coefficient, Cp, plotted against dimensionless axial 

location, X/Cx for Z = 50, 75, and 97.5 percent respectively for Re = 343,000. The plots show cases with 

no jet, one jet at +30° and one jet at –30°. The plots do not show any noticeable differences due to the jet 

on the static pressure and accordingly on the blade loading. The effect on the leakage and the total 

pressure losses are examined later. 

Figures 8(a) to (c) show the static pressure coefficient plotted against dimensionless axial location, 

X/Cx for Z = 50, 75, and 97.5 percent span respectively for Re = 343,000. On the plots shown, cases with 

no Jet, three jets with 30q plus and 30q minus and 11 jets with 30q plus. 

In this section our study will focus of the following cases: (1) No jet, (2) One jet, minus 30q, (3) One 

jet, 90q, (4) Three jets, plus 30q, and (5) Eleven jets, plus 30q. 

Figure 9(a) shows the velocity vectors colored by the velocity magnitude, in a plane normal to Z and 

mid-gap, Re = 343,000, for no jets. The plot shows the flow velocity is very high in the tip gap and in the 

aft 30 percent of the blade chord. Figure 9(b) shows the velocity vectors colored by the velocity 

magnitude, in a plane normal to Z and mid-gap, Re = 343,000, for one jet, minus 30q. The plot shows 

how the flow is accelerated due to the presence of the jet. This fact is responsible for having more leakage 

(as will be shown later). Figure 9(c) shows the velocity vectors colored by the velocity magnitude, in a 
NASA/TM—2014-218494 6 

plane normal to Z and mid-gap, Re = 343,000, for one jet, 90q. The plot shows how the flow is 

decelerated due to the presence of the jet. However the flow is accelerated upstream and around the jet. 

This fact is responsible for having more leakage (as will be shown later). Figures 9(d) and (e) show the 

velocity vectors colored by the velocity magnitude, in a plane normal to Z and mid-gap, Re = 343,000, for 

(3) jets, plus 30q and (11) jets, plus 30q, respectively. The plots show how the flow is decelerated due to 

the presence of the jets near the leading edge to 30 percent chord region while it is accelerated in the aft 

region of the tip. 

Figure 10(a) shows the velocity magnitude plotted versus X/Cx at two lines: mid-distance between 

the blade tip and the casing, one on the pressure side (PS) and one on the suction side (SS), Re = 343,000, 

no jets and one jet, minus 30q. The vertical line shows the location of the jet. For the case of no jet, the 

velocity magnitude on SS side is higher that than on PS except in the range of X/Cx from 0.38 to 0.48. In 

the case of one jet, the blowing ratio (jet outlet velocity/inlet velocity) is about 2. The velocity magnitude 

on PS side did not change much. However on the SS side the signature of the jet is very clear with a slight 

decrease in the velocity magnitude. 

Figure 10(b) shows the velocity magnitude plotted versus X/Cx at two lines: mid-distance between 

the blade tip and the casing, one on the pressure side (PS) and one on the suction side (SS), Re = 343,000, 

no jets and one jet, 90q. The vertical line shows the location of the jet. For the case of no jet, the velocity 

magnitude on SS side is higher that than on PS except in the range of X/Cx from 0.38 to 0.48. In the case 

of one jet, the blowing ratio (jet outlet velocity/inlet velocity) is about 2. The velocity magnitude on PS 

side did not change much. However on the SS side the signature of the jet is very clear with a decrease in 

the velocity magnitude down to 50 percent. 

Figure 10(c) shows the velocity magnitude plotted versus X/Cx at two lines: mid-distance between the 

blade tip and the casing, one on the pressure side (PS) and one on the suction side (SS), Re = 343,000, no 

jets and (3) jets, plus 30q. The vertical lines show the location of the jets. In the case of (3) jets the blowing 

ratio (ratio of jet exit velocity to the inlet) of the jets (starting from the upstream) are: 1.14, 1.31, and 1.48, 

respectively (see Fig. 11). The velocity magnitude on PS side changed slightly. However on the SS side the 

signature of the (3) jets is very clear with decrease in the velocity magnitude down to close to zero. 

Figure 10(d) shows the velocity magnitude plotted versus X/Cx at two lines: mid-distance between 

the blade tip and the casing, one on the pressure side (PS) and one on the suction side (SS), Re = 343,000, 

no jets and (11) jets, plus 30q. In the case of (11) jets the blowing ratio of the jets (starting from the 

upstream) are: 0.79, 1.13, 1.2, 1.85, 2.03, 2.11, 2.16, 2.14, 2.2, 2.2, and 2.2, respectively. The velocity 

magnitude on the pressure side changed considerably by almost a factor of 2 reduction in the presence of 

the jets. On the suction side the signature of the (11) jets is very clear with considerable decrease in the 

velocity magnitude. 

Figure 11 shows the blowing ratio for the individual holes for the 3-jets and 11-jets cases. For the 

same coolant plenum pressure ratio, the blowing ratio rises from 1 near the leading edge to almost 2 near 

30%Cx and stays close to 2 for the remainder of the holes. 

Figure 12(a) to (e) shows the total pressure contours at a plane 1.086 Cx from the leading edge for: 

12(a) No jet, 12(b) One jet, minus 30q, 12(c) One jet, 90q, 12(d) Three jets, plus 30q, and 12(e) Eleven 

jets, plus 30q, respectively. The loss core in the leakage vortex is clearly visible as is the wake and 

passage vortex. Figure 12(d) shows the total pressure for the 3 hole, plus 30 case. there is a clear 

reduction in the loss core and a shift of the leakage vortex away from the suction side. An area average 

(see Fig. 16) was taken for all cases and will be compared later. 

Figure 13(a) to (e) shows the static pressure contours at the blade tip and the suction side for: 12(a) No 

jet, 12(b) One jet, minus 30q, 12(c) One jet, 90q, 12(d) Three jets, plus 30q, and 12(e) Eleven jets, plus 30q, 

respectively. The areas that show low pressure (either on the tip or the suction side) indicate a momentum 

transfer from other regions and hence enhance the vortex structure and increase the losses. The data 

correlates very well with loss coefficient, area-averaged shown in Figure 16 (to be discussed below). 

 
NASA/TM—2014-218494 7 

Figure 14 shows the tip leakage mass flow as percent of the total inlet mass flow for (6) segments 

along the suction side of the gap, Re = 343,000, no jets, one jet, minus 30q, one jet 90q, three jets, plus 

30q, and eleven jets, plus 30q. It is clear that a majority of the leakage flow exits the tip gap in the aft 

30 percent of the blade chord. With one jet (minus 30q) there is noticeable increase in the leakage in all 

segments (as expected) since the jet is flowing in the same direction of the leaking flow. With one jet 

(90q) there is noticeable increase in the leakage in segments (1) and (4). In the case of (3) jets, plus 30q a 

noticeable reduction in the leakage is seen in segment (4). With (11) jets, plus 30q a reduction in the 

leakage is seen in all segments with a notable one occurring in segment (4). The largest impact seems to 

be from having a jet either at 90q to the tip surface or blowing toward the pressure side at 30q. There is a 

shift in leakage distribution from 42 to 58 percent chord. This makes sense because the presence of one 

jet, while able to perhaps cause a loss in momentum of the leakage flow is unable to block or impact the 

leakage vortex. The leakage flow merely circumvents the jet and shifts to an aft location. 

Figure 15 shows the total tip leakage mass flow as percent of the No jet case, Re = 343,000, one jet, 

minus 30q (with 4 percent increase), one jet 90q (with 4 percent decrease), three jets, plus 30q (with 

5 percent decrease), and eleven jets, plus 30q (with 12 percent decrease). These results are consistent with 

what is discussed in Figure 10. 

Figure 16 show loss coefficient, area averaged at 1.086 Cx as percent of the no jet case, 

Re = 343,000, one jet, minus 30q (with 24 percent decrease), one jet 90q (with 33 percent decrease), 

three jets, plus 30q (18 percent decrease), and eleven jets, plus 30q (with 5 percent increase). 

The low blowing ratio jets near the leading edge seem to redirect the flow in the tip toward the trailing 

edge and delay the mixing of the tip leakage with the suction side flow thus reducing the pressure loss 

coefficient. The high blowing ratio and strengthening of the loss core for the 11 jet case leads to an increase 

in losses. It is possible that individual plena with different pressure ratios for the aft 70 percent and for the 

leading edge holes would be beneficial to loss reduction. Table 3 shows a summary of all cases. 

Accomplishing both reductions in the tip leakage flow and the loss coefficient, area averaged would 

require strategic allocation of: (1) number of jets, (2) jets location, and (3) jet angle. Blowing ratio is 

expected to play a role but was not included in this study. 

Conclusions 

The following conclusions are drawn from this investigation: 

1. The turbulence model selected (SST) and the grid topology designed showed good agreement in 

predicting pressure distribution along the EEE blade for a wide range of Re (from 85,000 to 685,000), 

2. The study was focused on Re = 343,000 and examining different jet configurations (one, three and 

eleven) distributed on the camber cord. Different jet angles were examined; 90q, 30q plus and 

30q minus. 

3. With no jet about 50 percent of the mass leakage took place in the later 30 percent of the blade near 

the trailing edge. 

4. Jets with 30q minus increase the mass flow while showing a decrease in loss coefficient, area 

averaged. This could be attributed to the fact that jets with 30 minus will be in the same direction 

where the flow leakage occur thus increase the mass flow leaking. However, these jets will be able to 

break the tip leakage vortex and reducing its strength and thus reduce the losses. 

5. On the contrary to the above, 30q plus will reduce the leakage but will increase the losses. 

6. A future exercise is to use some jets with 30 plus and others with 30 minus to achieve both leakage 

reduction as well as loss coefficient reduction. 
NASA/TM—2014-218494 8 

TABLE 1.—CASCADE PARAMETERS 

Axial 

chord, 

Cx

[mm] 

Pitch, 

LI

[mm] 

Span 

[mm] 

Inlet 

flow 

angle 

Exit 

flow 

angle 

132.59 132.59 152.4 +29.7q –65q

TABLE 2.—PARAMETERS OF THE 

JETS CONFIGURATIONS 

Case 1 Jet 3 Jets 11 Jets 

Chordwise 

location (%Cx) 

23 23 to 69 

(equally 

spaced) 

13 to 90 

(equally 

spaced) 

Holes diameter 

(mm) 

5 5 5 

Length to 

diameter (L/D) 

ratio 

1.5 1.5 1.5 

Plenum 

pressure (kPa) 

36.942 36.942 36.942 

Plenum 

temperature 

(K) 

300 300 300 

TABLE 3.—SUMAMRY OF LEAKAGE AND 

LOSSES FOR ALL CASES 

Case Tip leakage 

mass flow 

Loss coefficient,

area averaged 

No Jets 100 100 

One Jet, Minus 30q 104 76 

One Jet, 90q 96 67 

(3) Jets, Plus 30q 95 82 

(11) Jets, Plus 30q 88 105 

 
NASA/TM—2014-218494 9 

Figure 1.—Computation domain and boundary conditions. 

Figure 2.—Cascade Layout with EEE Blades at E1 = 38.8q (I = +9.1q). 

 
NASA/TM—2014-218494 10 

Figure 3(a).—Cps comparison, Exp. and CFD, 

Re = 85,000, Z = 50%, No Gap. 

Figure 3(b).—Cps comparison, Exp. and CFD, 

Re = 85,000, Z = 75%, No Gap. 

Figure 3(c).—Cps comparison Exp. and CFD, 

Re = 85,000, Z = 97.5%, No Gap. 

Figure 3(d).—Cp-total at 1.086 Cx, comparison, Exp. and 

CFD, Re = 85,000, Z = 50%, No Gap. 

 

-1.00

-0.50

0.00

0.50

1.00

1.50

0.00 0.20 0.40 0.60 0.80 1.00

Cps

X/Cx

Exp.-Re-85k

RSM

SST

-1.00

-0.50

0.00

0.50

1.00

1.50

0.00 0.20 0.40 0.60 0.80 1.00

Cps

X/Cx

Exp.-Re-85k

RSM

SST

-1

-0.5

0

0.5

1

1.5

0 0.2 0.4 0.6 0.8 1

Cps

X/Cx

RSM

SST

-0.1

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

-0.5 -0.4 -0.3 -0.2 -0.1 0 0.1 0.2 0.3 0.4 0.5

Cp-Tot.

Y/Pitch

Exp.-Re-85k

RSM

SST
NASA/TM—2014-218494 11 

Figure 4(a).—Cps comparison, Exp. and CFD, 

Re = 343,000, Z = 50%, No Gap. Figure 4(b).—Cps comparison, Exp. and CFD, 

Re = 343,000, Z = 75%, No Gap. 

Figure 4(c).—Cps comparison, Exp. and CFD, 

Re = 343,000, Z = 97.5%, No Gap. 

Figure 4(d).—Cp-total at 1.086 Cx, comparison, Exp. and 

CFD, Re = 343,000, Z = 50%, No Gap. 

 

-1

-0.5

0

0.5

1

1.5

0 0.2 0.4 0.6 0.8 1

Cps

X/Cx

Exp.-Re-343k

RSM

SST

-1

-0.5

0

0.5

1

1.5

0 0.2 0.4 0.6 0.8 1

Cps

X/Cx

Exp.-Re-343k

RSM

SST

-1

-0.5

0

0.5

1

1.5

0 0.2 0.4 0.6 0.8 1

Cps

X/Cx

RSM

SST

-0.1

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

-0.5 -0.4 -0.3 -0.2 -0.1 0 0.1 0.2 0.3 0.4 0.5

Cp-Tot.

Y/Pitch

Exp.Re-343k

RSM

SST
NASA/TM—2014-218494 12 

Figure 5(a).—Cps comparison, Exp. and CFD, 

Re = 685,000, Z = 50%, No Gap. Figure 5(b).—Cps comparison, Exp. and CFD, 

Re = 685,000, Z = 75%, No Gap. 

Figure 5(c).—Cps comparison, Exp. and CFD, Re = 

685,000, Z = 97.5%, No Gap. 

Figure 5(d).—Cp-total at 1.086 Cx, comparison, Exp. and 

CFD, Re = 685,000, Z = 50%, No Gap. 

Figure 6.—Loss Coefficient, Area Averaged, for all cases 

examined, versus Re and RSM and SST turbulence 

models. 

 

-1

-0.5

0

0.5

1

1.5

0 0.2 0.4 0.6 0.8 1

Cps

X/Cx

Exp.-Re-685k

RSM

SST

-1

-0.5

0

0.5

1

1.5

0 0.2 0.4 0.6 0.8 1

Cps

X/Cx

Exp.-Re-685k

RSM

-1

-0.5

0

0.5

1

1.5

0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1

Cps

X/Cx

RSM

SST

-0.1

0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

-0.5 -0.4 -0.3 -0.2 -0.1 0 0.1 0.2 0.3 0.4 0.5

Cp-Total

Y/Pitch

Exp.-Re-685k

RSM

SST

0.04

0.06

0.08

0.1

0.12

0.14

0.16

80000 380000 680000

Loss Coefficient, Area 

Averaged

Re

Exp. Re = 85,000

Exp. Re = 343,000

Exp. Re = 685,000

CFD, SST Model

CFD, RSM Model
NASA/TM—2014-218494 13 

Figure 7(a).—Cps comparison, CFD, Re = 343,000, 

Z = 50%, No Jet, One Jet, Plus and Minus 30q. Figure 7(b).—Cps comparison, CFD, Re = 343,000, 

Z = 75%, No Jet, One Jet, Plus and Minus 30q. 

Figure 8(a).—Cps comparison, CFD, Re = 343,000, Z = 

50%, No Jet, 3 Jets, Plus and Minus 30 and 11 Jets 

Plus 30q. Figure 8(b).—Cps comparison, CFD, Re = 343,000, 

Z = 75%, No Jet, 3 Jets, Plus and Minus 30 and 

11 Jets Plus 30q. 

 

-1

-0.5

0

0.5

1

1.5

0 0.2 0.4 0.6 0.8 1

Cps

X/Cx

CFD-W-Gap-No-Jet Z=50%

CFD-W-Gap-Plus-30-Deg-Jet

CFD-W-Gap-Minus-30-Deg-Jet

-1

-0.5

0

0.5

1

1.5

0 0.2 0.4 0.6 0.8 1

Cps

X/Cx

CFD-W-Gap-No-Jet Z=75%

CFD-W-Gap-Plus-30-Deg-Jet

CFD-W-Gap-Minus-30-Deg-Jet

-1

-0.5

0

0.5

1

1.5

0 0.2 0.4 0.6 0.8 1

Cps

X/Cx

CFD-W-Gap-No-Jet Z=50%

CFD-W-Gap-3-Jets-Minus-30-Deg

CFD-W-Gap-3-Jets-Plus-30-Deg

CFD-W-Gap-11-Jets-Plus-30-Deg

-1

-0.5

0

0.5

1

1.5

0 0.2 0.4 0.6 0.8 1

Cps

X/Cx

CFD-W-Gap-No-Jet Z=75%

CFD-W-Gap-3-Jets-Minus-30-Deg

CFD-W-Gap-3-Jets-Plus-30-Deg

CFD-W-Gap-11-Jets-Plus-30-Deg
NASA/TM—2014-218494 14 

Figure 9(a).—Velocity vectors colored by the velocity 

magnitude, in a plane normal to Z and mid-gap, 

Re = 343,000, No Jet. 

 

Figure 9(b).—Velocity vectors colored by the velocity 

magnitude, in a plane normal to Z and mid-gap, 

Re = 343,000, One Jet, Minus 30q. 

 

Figure 9(c).—Velocity vectors colored by the velocity 

magnitude, in a plane normal to Z and mid-gap, 

Re = 343,000, One Jet, 90q. 

 

Figure 9(d).—Velocity vectors colored by the velocity 

magnitude, in a plane normal to Z and mid-gap, 

Re = 343,000, Three Jets, Plus 30q. 

Figure 9(e).—Velocity vectors colored by the velocity 

magnitude, in a plane normal to Z and mid-gap, 

Re = 343,000, Eleven Jets, Plus 30q. 

 

 
NASA/TM—2014-218494 15 

Figure 10(a).—Velocity magnitude versus X/Cx at PS 

and SS for No jet and One Jet, Minus 30q. 

Figure 10(b).—Velocity magnitude versus X/Cx at PS and 

SS for No jet and One Jet, 90q. 

Figure 10(c).—Velocity magnitude versus X/Cx at PS 

and SS for No jet and Three Jets, Plus 30q. Figure 10(d).—Velocity magnitude versus X/Cx at PS and 

SS for No jet and Eleven Jets, Plus 30q. 

Figure 11.—Jets velocity magnitude/ Inlet velocity 

versus X/Cx for Three and Eleven Jets, Plus 30q. 
NASA/TM—2014-218494 16

Figure 12(a).—Contours of the total pressure at 

1.086Cx (from leading edge), for No jet(s). 

Figure 12(b).—Contours of the total pressure at 1.086Cx 

(from leading edge), for One Jet, Minus 30q. 

Figure 12(c).—Contours of the total pressure at 

1.086Cx (from leading edge), for One Jet, 90q. 

Figure 12(d).—Contours of the total pressure at 1.086Cx 

(from leading edge), for Three Jets, Minus 30q. 

Figure 12(e).—Contours of the total pressure at 

1.086Cx (from leading edge), for Eleven Jets, 

Minus 30q. 
NASA/TM—2014-218494 17

Figure 13(a).—Contours of the static pressure at the 

blade top and SS, for No jet(s). 

Figure 13(b).—Contours of the static pressure at the blade 

top and SS, for One jet, Minus 30q. 

Figure 13(c).—Contours of the static pressure at 

the blade top and SS, for One jet, 90q. 

Figure 13(d).—Contours of the static pressure at the blade 

top and SS, for Three jets, Plus 30q. 

Figure 13(e).—Contours of the static pressure at the 

blade top and SS, for Eleven jets, Plus 30q. 
NASA/TM—2014-218494 18 

Figure 14.—Tip leakage mass flow as percent of total inlet mass flow for (6) segments 

along the suction side of the gap, Re = 343,000, No Jets, One Jet Minus 30q, One Jet 

90q, Three Jets, Plus 30q, and Eleven Jets, Plus 30q. 

Figure 15.—Total tip leakage mass flow as percent of 

the No Jet case, Re = 343,000, No Jets, One Jet 

Minus 30q, One Jet 90q, Three Jets, Plus 30q, and 

Eleven Jets, Plus 30q. 

Figure 16.—Loss coefficient, area averaged at 

1.086 Cx as percent of the No Jet case, Re = 

343,000, No Jets, One Jet Minus 30q, One Jet 90q, 

Three Jets, Plus 30q, and Eleven Jets, Plus 30q. 
NASA/TM—2014-218494 19 

References 

1. Bunker, R.S., “A Review of Turbine Blade Tip Heat Transfer,” Annals of the New York Academy 

of Science, Vol. 934, pp 64-79, January 2006. 

2. Glezer, B., Harvey, N.W., Camci, C., Bunker, R.S., Ameri, A.A., “Turbine Blade Tip Design and 

Tip Clearance Treatment,” notes from von Karman Lecture Series 2004, June 2004. 

3. Zhou, C., and Hodson, H., “The Tip Leakage flow of an Unshrouded High Pressure Turbine Blade 

With Tip Cooling,” ASME-GT2009-59637, June 2009. 

4. Ameri, A.A., Steinthorsson, E., Rigby, D.L., “Effect of Squealer Tip on Rotor Heat Transfer and 

Efficiency,” Journal of Turbomachinery, Vol. 120, pp 753-759, October 1998. 

5. Ameri, A.A., Steinthorsson, E., Rigby, D.L., “Effects of Tip Clearance and Casing Recess on Heat 

Transfer and Stage Efficiency in Axial Turbines,” Journal of Turbomachinery, Vol. 121, pp 683-

693, October 1999. 

6. Hofer, T., Arts, T., “Aerodynamic Investigation of the Tip Leakage Flow for Blades With Different 

Tip Squealer Geometries at Transonic Conditions,” ASMEGT2009-59909, June 2009. 

7. Krishnababu, S.K., Dawes, W.N., Hodson, H.P., Lock, G.D., Hannis, J., Whitney, C., “AeroThermal Investigations of Tip Leakage Flow in Axial Flow Turbines, Part II -Effect of Relative 

Casing Motion,” ASME 2007-GT-27957, May 2007. 

8. Shyam V., Ameri A., Chen J.P., “Analysis of Unsteady Tip and Endwall Heat Transfer in a Highly 

Loaded Transonic Turbine Stage,” Proceedings of the ASME TURBO Expo 2010, Glasgow, June 

2010, J. Turbomach. 134, 041022 (2012). 

9. Shyam, V., Ameri, A., “Comparison of Various Supersonic Turbine Tip Designs to Minimize 

Aerodynamic Loss and Tip Heating.” NASA/TM—2012-217607. 

10. Wheeler, A., Atkins, N., He, L., “Turbine Blade Tip Heat Tansfer in Low Speed and High Speed 

Flows,” J.Turbomachinery, Vol. 133, No. 4 2011. 

11. N. M. Rao and C. Camci, “Axial turbine tip desensitization by injection from a tip trench—part 1: 

effect of injection mass flow rate,” in Proceedings of the ASME Turbo Expo: Power for Land, Sea 

and Air Conference, pp. 1075–1088, 2004, ASME paper no. GT2004-53256. 

12. Couch, E., Christophel, J., Hohlfeld, E., Thole, K., “Comparison of Measurements and Predictions 

for Blowing from a Turbine Blade Tip,” Journal of Propulsion and Power, Vol. 21, No.2, MarchApril 2005. 

13. Bogard, D., and Thole, K., 2006, “Gas Turbine Film Cooling,” Journal of Propulsion and Power, 

22(2), 249 270, DOI:10.2514/1.18034. 

14. Canelli, C., Castelanelli, M., Luccoli, R., Abba, L., Corcoruto, S., Traverso, S., “Numerical 

Prediction of Film Cooling on the Squealer Tip of Gas Turbine Blade,” 59th Congresso Annuale 

ATI – Geneva 2004. 

15. Ibrahim, M.B., Kartuzova, O. and Volino, R.J., “Experimental and Computational Investigations of 

Separation and Transition on a Highly Loaded Low Pressure Turbine Airfoil: Part 1 – Low 

Freestream Turbulence Intensity,” ASME Paper IMECE2008-68879, 2008. 

16. Ibrahim, M.B., Kartuzova, O., Doucet, D.J. and Volino, R.J. “Les Flow Control Simulations For 

Highly Loaded Low Pressure Turbine Airfoil (L1a) Using Pulsed Vortex Generator Jets,” ASME 

Paper GT2010-23015, 2010. 

17. McVetta, A., Giel, P., Welch, G., “Aerodynamic Investigation of Incidence Angle Effects in a Large 

Scale Transonic Turbine Cascade,” AIAA 2012-3879, 48th AIAA/ASME/SAE/ASEE Joint 

Propulsion Conference & Exhibit. July. 

18. ANSYS FLUENT Documentation, 2009. 
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Computational Investigation of Novel Tip Leakage Mitigation Methods for High

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Ibrahim, Mounir; Gupta, Abhinav; Shyam, Vikram

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13. SUPPLEMENTARY NOTES

14. ABSTRACT

This paper presents preliminary findings on a possible approach to reducing tip leakage losses. In this paper a computational study was conducted on the Energy Efficient

Engine (EEE) High Pressure Turbine (HPT) rotor tip geometry using the commercial numerical solver ANSYS FLUENT. The flow solver was validated against aerodynamic

data acquired in the NASA Transonic Turbine Blade Cascade facility. The scope of the ongoing study is to computationally investigate how the tip leakage and overall blade

losses are affected by (1) injection from the tip near the pressure side, (2) injection from the tip surface at the camber line, and (3) injection from the tip surface into the tip

separation bubble. The objective is to identify the locations on the tip surface at which to place appropriately configured blowing keeping in mind the film cooling application

of tip blowing holes. The validation was conducted at Reynolds numbers of 85,000, 343,000, and 685,000 and at engine realistic flow conditions. The coolant injection

simulations were conducted at a Reynolds number of 343,000 based on blade chord and inlet velocity and utilized the SST turbulence model in FLUENT. The key parameters

examined are the number of jets, jet angle and jet location. A coolant to inlet pressure ratio of 1.0 was studied for angles of +30°, -30°, and 90° to the local free stream on the

tip. For the 3 hole configuration, 3 holes spaced 3 hole diameters apart with length to diameter ratio of 1.5 were used. A simulation including 11 holes along the entire mean

camber line is also presented (30° toward suction side). In addition, the effect of a single hole is also compared to a flat tip with no injection. The results provide insight into tip

flow control methods and can be used to guide further investigation into tip flow control. As noted in past research it is concluded that reducing leakage flow is not necessarily

synonymous with reducing losses due to leakage.

15. SUBJECT TERMS

Leakage; Blowing; Supersonic; Turbine; Engine; Camber; Hole geometry; Turbulence models; Aerodynamics

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