Awatef Hamed, Widen Tabakoff, Rohan Swar, Dongyun Shin, and Nathanial Woggon

University of Cincinnati, Cincinnati, Ohio

Robert Miller

Glenn Research Center, Cleveland, Ohio

Combined Experimental and Numerical Simulations 

of Thermal Barrier Coated Turbine Blades Erosion

NASA/TM—2013-217857

April 2013
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Awatef Hamed, Widen Tabakoff, Rohan Swar, Dongyun Shin, and Nathanial Woggon

University of Cincinnati, Cincinnati, Ohio

Robert Miller

Glenn Research Center, Cleveland, Ohio

Combined Experimental and Numerical Simulations 

of Thermal Barrier Coated Turbine Blades Erosion

NASA/TM—2013-217857

April 2013

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135
Acknowledgments

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at the NASA Glenn Research Center.

Level of Review: This material has been technically reviewed by technical management. 
NASA/TM—2013-217857 1 

Combined Experimental and Numerical Simulations of

Thermal Barrier Coated Turbine Blades Erosion

Awatef Hamed, Widen Tabakoff, Rohan Swar, Dongyun Shin, and Nathanial Woggon 

University of Cincinnati

Cincinnati, Ohio 45220 

Robert Miller

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135 

Abstract

A combined experimental and computational study was conducted to investigate the erosion of 

thermal barrier coated (TBC) blade surfaces by alumina particles ingestion in a single stage turbine. In the 

experimental investigation, tests of particle surface interactions were performed in specially designed 

tunnels to determine the erosion rates and particle restitution characteristics under different impact 

conditions. The experimental results show that the erosion rates increase with increased impingement 

angle, impact velocity and temperature. In the computational simulations, an Euler-Lagrangian two stage 

approach is used in obtaining numerical solutions to the three-dimensional compressible Reynolds

Averaged Navier-Stokes equations and the particles equations of motion in each blade passage reference 

frame. User defined functions (UDF) were developed to represent experimentally-based correlations for 

particle surface interaction models which were employed in the three-dimensional particle trajectory 

simulations to determine the particle rebound characteristics after each surface impact. The experimentally based erosion UDF model was used to predict the TBC erosion rates on the turbine blade surfaces 

based on the computed statistical data of the particles impact locations, velocities and angles relative to 

the blade surface. Computational results are presented for the predicted TBC blade erosion in a single

stage commercial APU turbine, for a NASA designed automotive turbine, and for the NASA turbine 

scaled for modern rotorcraft operating conditions. The erosion patterns in the turbines are discussed for 

uniform particle ingestion and for particle ingestion concentrated in the inner and outer 5 percent of the 

stator blade span representing the flow cooling the combustor liner. 

Introduction

Turbomachinery erosion presents a challenging problem when gas turbine engines operate in dusty 

environments (Refs. 1 to 3). Some of the mechanisms that cause particle ingestion are (a) the vortex from 

engine inlet-to-ground during high power setting with the aircraft standing or moving on the runway; (b) 

storms transporting sand to several thousand feet altitude; (c) thrust reverser afflux at low airplane speed 

blowing sand, ice and other particles into the engine inlets. Erosive solid particles may also be produced 

during the combustion process, from the burning of different types of heavy oils or synthetic fuels. 

Helicopter engines are especially susceptible to large amounts of dust and sand ingestion during hover, 

takeoff and landing.

It is very difficult to remove all solid particles from the gas stream without taxing the performance of 

gas turbine engines (Refs. 4 and 5). Even small particles of 1 to 30 m size have been known to be very 

damaging to the exposed components of coal burning turbines (Ref. 6). In turbomachinery, particle 

impacts are known to increase tip clearances and blade surface roughness and produce changes in the 

blade leading and trailing edges (Ref. 7). Because TBC erosion is detrimental to the thermal protection of 

turbine blades, it has been identified as a life limiting factor for gas turbines (Refs. 8 to 10). With the 
NASA/TM—2013-217857 2 

increasing use of thermal barrier technology to protect highly loaded and rotating turbine components 

(Ref. 11), further studies are needed to support thermal protection for life extension. 

Different process combinations such as cutting, fatigue, brittle fracture and melting mechanisms have 

been proposed to govern material removal by erosion. According to experimental studies of blade alloys 

and coating materials, their erosion rates are influenced by particle impact velocities and impingement 

angles and by the operating temperatures (Ref. 3). Experimental characterization of material erosion 

resistance requires special facilities that control particle-laden flow around the sample to achieve the 

desired impact conditions over the tested samples (Ref. 12). Prior TBC erosion test results have 

demonstrated that electron beam-physical vapor deposited (EB-PVD) TBC erosion rates are an order of 

magnitude less than plasma sprayed (PS) TBC and that both increase linearly with particles impact 

velocity (Refs. 13 and 14). Experimental studies of particle surface impacts are also necessary to provide 

particle rebound characteristics over the range of impinging conditions encountered in turbomachines 

(Refs. 3 and 7). 

Blade erosion in turbomachines is affected by rotational speed and flow conditions, blade geometry and 

blade row location, blade material and particle characteristics (Refs. 3 and 7). Under two-phase flow 

conditions, the gas and particles experience different degrees of turning through the blade passages. 

Deviation from the gas flow path increases with particle inertia causing repeated impacts with the various 

surfaces. Surface impacts alter the direction and velocity of the particles as well as their distribution through 

subsequent stages. Experimental studies that simulate erosive particle impact conditions in the engine 

environment are essential to blade and coating material development. The associated blade surfaces erosion 

requires knowledge of the particle three-dimensional trajectories and their impact statistics on the various 

engine surfaces in addition to surface material erosion behavior under the impact conditions.

Trajectory simulations are based on the numerical integration of the particles’ equations of motion 

through the successive turbomachinery stationary and rotating blade rows. The simulations require the 

three-dimensional flow field and blade passage geometries as inputs and a model for particle restitution 

conditions following each surface impact. The basis for particle trajectory simulations in turbomachines 

continues to be Eulerian-Lagrangian with one way coupling between the particles and flow (Refs. 15 and 

16). Currently three-dimensional flow field solutions of the Reynolds Averaged Navier-Stokes equations 

for turbulent flow through the blade passages (Ref. 16) are used in turbomachinery trajectory simulations. 

Hamed and Tabakoff (Ref. 17) developed a methodology to predict turbomachinery blade surface 

erosion using blade surface statistical impact data computed from particle trajectory simulations and 

correlations based on blade material erosion test results. It was used in the prediction of blade erosion in 

both axial and radial compressors and turbines (Ref. 3) and erosion of turbine blade coating developed for 

automotive and ground based gas turbine applications. 

In the current work, a combined experimental and numerical investigation was conducted to study the 

deterioration of thermal barrier coating on turbine blades by erosive particles. Experimental 

measurements were obtained for EB-PVD 7YSZ/PtAl coated Inconel 718 coupons in the high 

temperature erosion tunnel under different impact conditions by alumina particles. Measurements were 

also obtained for particle restitution characteristics at different impact angles relative to the coated sample 

using Particle Image Displacement Velocimetry (PIDV). The experimental measurements of surface 

erosion and particle restitution were used to develop empirical erosion and restitution models. The 

numerical simulations for the three-dimensional flow field and particle trajectories were combined with 

experimentally based surface interaction models to predict TBC erosion over the stator and rotor blade 

surfaces in a single stage commercial gas turbine and in a NASA automotive turbine design (Ref. 18) and 

scaled for modern rotorcraft operating conditions (Ref. 19).

Experimental Work

Erosion tests were carried out for 25.4 by 25.4 by 1.5875 mm (1 by 1 by 1/16 in.) Inconel 718 coupons 

coated with EB-PVD 7YSZ over a range of impact conditions by different size alumina particles. The 

coupons were mounted on a sample holder and placed at the designated angles in the high temperature 
NASA/TM—2013-217857 3 

Figure 1.—Schematic of erosion test facility. 

erosion wind tunnel shown schematically in Figure 1 and subjected to erosion by a calibrated mass of 

particles. The holder protected all but one target coupon surface that was exposed to particle impacts. The 

samples were weighed before and after the erosion tests to determine the weight loss due to erosion by the 

impacting particles.

The University of Cincinnati (UC) erosion wind tunnel (Ref. 12) shown schematically in Figure 1 

consists of the following components: particle feeder (A), main air supply pipe (B), combustor (C), 

particle pre-heater (D), particle injector (E), acceleration tunnel (F), test section (G) and exhaust tank (H). 

Abrasive particles of a given constituency and measured weight are placed into the particle feeder (A). 

The particles are fed into a secondary air source and blown into the particle preheater (D) and then to the 

injector (E), where they mix with the primary air supply (B), which is heated by the combustor (C). The 

particles are then accelerated via high velocity air in a constant-area steam-cooled duct (F) and impact the 

specimen in the test section (G). The particulate flow is then mixed with coolant and directed to the 

exhaust tank. As can be seen from Figure 1, the tunnel geometry is uninterrupted from the acceleration 

tunnel throughout the test section in order to preserve the aerodynamics of the flow passing over the 

samples. Particles’ impact velocities are calibrated with the tunnel air flow, while the particle 

impingement angle is controlled through the sample orientation. A measured mass of particles is fed into 

the flow after the test sample reaches equilibrium temperature. 

The erosion rate, ε in mg/g is defined as the ratio of the measured coated sample erosion mass loss, 

ΔW, to the mass of erosive particles impacting surface, Qp, which is determined from the tunnel 

calibration according to the sample size and impingement angle (Ref. 20). 

Qp

W =e (1) 
NASA/TM—2013-217857 4 

Particle restitution characteristics after impacting the coated coupons were measured in a second 

tunnel equipped with optical window in the test section for recording particle trajectories using high speed 

photographic methods at room temperature. A Phantom high speed digital camera (model V 9.1) with a 

Nikon lens was used to photograph the particle trajectories. The camera had a maximum frame speed of 

50,000 fps, but a frame rate of 27,000 fps (37 μs interval) and 256 256 resolution were appropriate for 

the light intensity obtained from a 250 mW Spectra-Physics laser source (454 to 676 nm wavelength). 

Referring to Figure 2, the laser sheet was 1.65 mm thick and 50 mm in height at the frame of interest and 

the exposure time could be set as low as 3 μs. The camera was connected to a computer and its graphical 

user interface (GUI) software was used to set the frame speed and resolution, after it was focused on a 

reference point in the laser sheet before recording the images. 

The magnitude and direction of the particle impact and rebound velocities were measured in the plane 

of the laser sheet. The x and y co-ordinates of a particle in each frame were determined using Phantom 

Cine Viewer image processing software. By comparing successive images, the trajectory of the particle 

was calculated using Particle Image Displacement Velocimetry (PIDV) technique (Refs. 21 to 23). The 

image of a 25.4 mm mark was used as a reference length to provide the actual distance the particle 

traveled between successive frames as shown in Figure 3. The particle velocity, V, was calculated based 

on the time between two successive frames:

r

FRd

t

S V = = (2) 

Figure 2.—Test section, laser sheet and camera.

Figure 3.—High speed photography images. 
NASA/TM—2013-217857 5 

Where:

δt time between successive frames, sec

F number of frames per second

V particle velocity

δS actual distance traveled by particle between two consecutive frames

R actual length of the reference mark in the test section

d distance traveled by particle as obtained from the pixel count 

r length of the reference mark as obtained from the pixel count

The velocity and directional restitution coefficients were determined from three successive frames of 

interest in order to exclude the post impact influence of the tunnel flow on the rebounding particle 

trajectory. Error estimations of the results obtained from high speed photography were carried out using 

the t-distribution method (Ref. 23) with 95 percent confidence interval.

Computational Work

Numerical simulations were conducted using ANSYS CFX 12 (Ref. 24) to calculate the threedimensional flow field and the associated 26 m particle trajectories in a single stage gas-turbine. The 

Eulerian-Lagrangian two stages, one-way particle interaction model approach was used because of the 

typically low particle concentration in gas turbines. 

Pointwise V16 (Ref. 25) code was used to generate the structured computational grids in the stator 

and rotor blade passages of the single stage turbines. Its elliptic solver was used to improve the grid 

quality especially in the high turning rotor blade passage. Figure 4 shows the three-dimensional H-O 

computational grids used in the NASA designed turbine flow field and particle dynamics simulations. 

Grid size was based on our prior experience in large commercial engines (Ref. 16). The grids were 

clustered near the blade passage surfaces to achieve the y+<3 at the first grid point next to the surface. 

The three-dimensional flow field solutions were obtained for the compressible Reynolds-Averaged 

Navier-Stokes equations in conservation form and the k-ω based Shear-Stress-Transport (SST) model. 

Advection fluxes were evaluated using the high-resolution scheme in ANSYS CFX 12 (Ref. 24) which is 

second-order accurate with convergence criteria of 10 to 5 rms residual. Because of the very high particle 

velocities in the turbine particularly after rebounding from rotor blade surface impacts, the turbulence 

stochastic effects are negligible and the aerodynamic forces on the particles are dominated by drag due to 

particle slip velocities relative to the gas velocities.

Figure 4.—Sample computational grids of the NASA designed turbine. 
NASA/TM—2013-217857 6 

The total pressure, total temperature inlet and flow direction were specified at the inlet boundary and 

uniform averaged static pressure was specified at the outlet boundary with the periodic interface boundary 

conditions invoked in the stator and rotor blade passages. A mixing plane model was used at the statorrotor interface. The simulations were conducted for particles entering the turbine in the same direction as 

the inlet gas flow but with 50 percent the value of the inlet gas velocities. 

The nominally 26 m erosive alumina particles were assumed to be ingested at the same temperature 

as the gas inflow. Convective heat transfer coefficient based on ANSYS CFX 12 (Ref. 24) correlation 

represented the interphase energy exchange. Particle trajectories were computed using forward integration 

and taking into consideration the momentum and heat exchange with the flow field based on the threedimensional flow solutions. The particle tracking integration time step was chosen locally such that the 

particles were advanced 100 time steps within each grid spacing used to obtain the three-dimensional 

turbine flow solutions (Ref. 24). 

Particle surface interaction modeling was based on experimental data obtained for TBC coated 

samples over a range of impact conditions by the alumina particles. The restitution and erosion prediction 

empirical models were implemented in ANSYS CFX 12 (Ref. 24) as user-defined functions. 

Results and Discussions

Experimental Results

Figure 5 presents sample results for the experimentally measured erosion rates variation with particle 

impingement angles for 10 mil EB-PVD 7YSZ coated Inconel 718 samples and for samples of the 

Inconel 718 substrate at 1093 °C (2000 °F) temperature and 366 m/s (1200 ft/s) impact velocity. The 

maximum erosion rate for the EB-PVD TBC coating occurred at 90° impingement angle while that for 

INCO substrate occurred between 20° to 30° which are typical brittle and ductile materials erosion 

behavior, respectively. One can see from Figure 6(a) that the nominally 26 m, mass mean erosive 

alumina particles have angular shapes with sharp corners. The particles size distribution is shown in 

Figure 6(b).

Figure 5.—Measured erosion rates due to 26 μm alumina particle 

impacts on TBC coated and uncoated samples (T = 1093 °C 

(2000 °F), V = 365.8 m/s (1200 ft/s)). 
NASA/TM—2013-217857 7 

 

(a) Scanning electron microscope image of sample particles. 

(b) Particle size distribution by weight. 

Figure 6.—Nominally 26 μm alumina particles. 

Figures 7 and 8 present experimental results for EB-PVD TBC erosion rates variation with 

impingement angles at 982 °C (1800 °F) and 871 °C (1600 °F), respectively. According to these results 

the erosion rates increase with temperatures and with particle impact velocities at all impingement angles. 

This is consistent with the TBC erosion test results presented at 90° impact angle in prior investigations 

by Tabakoff (Ref. 13) and Nicholls et al. (Ref. 14). The following experimentally based TBC erosion rate 

correlations were implemented in the UDF and used in ANSYS CFX 12 (Ref. 24) to compute the TBC 

erosion rates on the blade surfaces based on the local gas temperature T, particle impact velocity V1 and 

impingement angle 1, relative to the coated blade surface. 

( )( )n V1

3

1

2

1 1 b+b-b=e 28.3473.0301.1293.1 (3) 

Where:

n += 000476.00584.0 T

0

5

10

15

20

25

30

35

40

% by Weight

Particle Size
NASA/TM—2013-217857 8 

Figure 7.—TBC erosion rates by 26 μm alumina particles at 982 °C (1800 °F). 

Figure 8.—TBC erosion rates by 26 μm alumina particles at 871 °C (1600 °F). 
NASA/TM—2013-217857 9 

(a) Pre-erosion test ( 100 magnification)

 (b) Post-erosion test ( 1000 magnification)

Figure 9.—Sample TBC surface scans with electron microscope. 

Several post erosion tested samples were examined using Scanning Electron Microscope to 

investigate post erosion particle deposition. Sample results are presented in Figure 9 for pre and post 

erosion test samples. Particle deposition was rarely observed on the post erosion test samples, as shown in 

the post erosion 1000 magnification image. 

The experimental results were used to derive the particle velocity and angle restitution coefficients ev

and e relative to the impacted turbine coated surfaces as a function of the particles impingement angle 1. 

4

1

3

1

2

1 1

1

2 8218.01881.30075.49369.15.0 b-b+b-b+==

v

v

ev (4) 

4

1

3

1

2

1 1

1

2 e = 9569.04188.35837.3474.09832.0 b-b+b-b+= (5) 

However, ANSYS CFX 12 (Ref. 24) requires user defined functions in the form of restitution 

coefficients that are normal and tangential to the impacted surface. The experimental correlations of 

Equations (5) and (6) were therefore used to compute the normal and tangential restitution coefficients en

and et that are presented in Figure 10 using the following equations: 

( )

( ) 1

1

Sin

Sin = e

ee vn (6) 

( )

( ) 1

1

Cos

Cos = e

ee vt (7) 
NASA/TM—2013-217857 10

Figure 10.—Particle tangential and normal restitution coefficients 

variation with impact angles. 

Numerical Predictions of Turbine Blades TBC Erosion Based on Computed ThreeDimensional Flow Fields and Suspended Particle Trajectories Incorporating 

Experimentally Based Particle Surface Interaction Model

Three sets of numerical predictions of three-dimensional turbine flow fields, particle trajectories 

turbine blades TBC erosion are presented for (1) commercial APU turbine (2) NASA designed 

automotive turbine (Ref. 18) (3) a scaled version of NASA designed turbine for modern rotorcraft 

operating conditions (Ref. 19).

The particle material property used was 99 percent aluminum oxide (alumina). A density of 3.96 g/cc and 

a molecular weight of 101.961 g/mol described the particles’ mass. All particles were 26 m in diameter with 

an initial velocity in the axial direction of 64 percent of the gas flow velocity at the inlet plane.

Commercial APU Turbine

Figure 11 presents carpet plots of the computed absolute Mach number at mid-span through the stator 

and the relative Mach number through the rotor, while Figure 12 presents the corresponding static 

pressure. The blade wakes are visible in Figure 11 and also a strong expansion wave at the rotor pressure 

surface trailing edge that extends across the blade passage towards the opposite blade suction surface is 

visible in both Figures 11 and 12. The two figures also indicate a second shock emanates off the rotor 

trailing edge suction surface and propagates downstream.

Figure 13 shows 25 sample trajectories with color contouring of their velocity magnitude for 26 m 

alumina particles seeded uniformly across the turbine inlet. The figure indicates that after the particles 

enter the stator blade row flow field they continue along their axial path until they impact the stator blade 

surface. Mainly at the pressure surface and the leading edge of the suction surface. The particles then

enter the rotor blade passage and impact the suction surface of the rotor blade because of their lower

absolute velocities compared to the gas flow as they leave the stator and enter the rotor as demonstrated 

by their absolute and relative velocities shown schematically in Figure 14. The particles rebound with 

high circumferential velocities after their rotor blade suction surface impacts that cause them to travel 

radially outward to the shroud. Many particles are seen to reenter the stator blade passage and impact
NASA/TM—2013-217857 11

the aft portion of the stator blades suction surface with the high velocities they acquired from the rotor blade 

surface impacts. Some particles reenter the stator blade passage only momentarily then are dragged by the 

flow back into the rotor blade passage. Eventually the majority of these particles continue to travel down the 

rotor blade passage impacting the pressure surface of the turbine rotor at a significantly lower velocity.

Figure 11.—Mach number contours at the 50 percent

blade span (Commercial APU Turbine). 

Figure 12.—Computed normalized pressure near mid-span 

(Commercial APU Turbine). 
NASA/TM—2013-217857 12

Figure 13.—Sample three-dimensional particle trajectories in the commercial APU turbine.

Figure 14.—Schematic showing effect of particle slip velocity at the 

stator exit on entrance condition in the rotor reference frame. 

Figure 15 presents the predicted thermal barrier coating erosion rates on the stator and rotor blades of 

the commercial APU turbine. The figures indicate relatively low uniform thermal barrier coating erosion 

rate on the aft half of the stator blade pressure surface and high erosion rates on the stator blade suction 

surface at the corner of the tip and trailing edge. This high erosion rate on the stator suction surface is 

caused by the high velocity particles that reenter the stator blade row after impacting the rotor suction 

surface because they lag the gas flow as they leave the stator blade row (Fig. 14). These same particles 

produce the highest erosion rate of the rotor blade suction surface at the shroud and leading edge corner. 

The particles are concentrated near the shroud by the centrifugal forces after they impact the rotor suction 

surface and gain the highest relative velocity. Several of the particles have second impacts as they reenter 

the rotor blade passage near the tip after they rebound from the stator suction surface impacts.

Figure 16 presents the computed results for particle trajectories when the particles are seeded only 

within 5 percent of the stator span at the hub and shroud. This seeding represents the suspended particle in 

the cooling secondary flow. Trajectory trends are similar to the previous case. The only obvious 

difference is the lack of particle impacts in the middle of the stator span due to the seeding locations. 

However, the particles seeded near the hub are still immediately thrown to the shroud when they impact 

the rotor blade suction surface. As the particles travel through the rotor they become concentrated at the 

shroud just as in the previous case.
NASA/TM—2013-217857 13

Figure 15.—TBC erosion rates on blade surfaces in the 

commercial APU turbine for particle ingested at 50 percent 

the gas velocity with uniform radial loading. 

Figure 16.—Computed three-dimensional particle trajectories for 

particle seeding within 5 percent of the rotor span and using the 

experimentally based UDF restitution model.
NASA/TM—2013-217857 14

Figure 17.—TBC erosion rates on blade surfaces in the 

commercial turbine for particle seeding within 5 percent 

of the rotor span at 50 percent of the gas velocity. 

TABLE 1.—COMPUTED OVERALL BLADES AND MAXIMUM LOCAL TBC EROSION RATES 

ON THE COMMERCIAL APU TURBINE 

Turbine Particle radial 

loading% span

Maximum local erosion rate, 

mg/g/cm2

Overall blade erosion rate,

mg/g

Stator Rotor

APU commercial turbine 5% 72.94 (Rotor suction side) 0.78 0.93

100% 31.16 (Stator suction side) 0.83 0.87

Figure 17 shows the corresponding thermal barrier coating erosion rates on the stator blade surfaces 

when the particle are seeded only within 5 percent of the stator span at the hub and shroud. Erosion on the 

stator pressure surface is seen to occur almost exclusively along these thin bands. Again the highest 

erosion rate is at the shroud and trailing edge corner of the stator blade suction surface that is caused by

the particles reentering the stator near the tip with higher velocities after impacting the rotor suction 

surface. The dominance of the centrifugal forces on the particles after their impacts with the rotor is 

evident even in the trajectories of the particles that were originally concentrated at the hub. 

A list of the predicted local maximum as well as the overall TBC erosion rates in the APU 

commercial turbine stator and rotor blades for both uniform and non-uniform particle injection is 

presented in Table 1.

In summary, the highest thermal barrier coating erosion rate in the APU turbine is always predicted 

on the rotor blade surface. This occurs mostly towards the shroud near the leading edge on the rotor 

suction surface. On the other hand, the pressure surface of the rotor always suffers the lowest thermal 

barrier coating erosion rate. Numerical simulations of particle trajectories were also conducted with a 

uniform restitution model based on the average value of the experimental measurements (Ref. 26). 

NASA Designed Original and Scaled Turbine for Modern Rotorcraft Operating Conditions

Numerical predictions of the three-dimensional flow field, particle trajectories and TBC blade coating 

erosion were conducted for a NASA designed automotive turbine whose geometry and operational 

characteristics are in the public domain as listed in Table 2 (Ref. 18). The original turbine operational 

conditions were scaled to meet the requirements of a modern rotorcraft engine comparable to Bell Ranger 

206 (Ref. 19). As shown in Table 3, the mass flow rate and rotating speed of the scaled turbine were 

based on the use of the following turbomachinery scaling laws: 
NASA/TM—2013-217857 15

( ) constant

RT

ND

;constant

RT

01

2

01

01 = = Dp

m& (8) 

TABLE 2.—NASA DESIGNED

AUTOMOTIVE TURBINE 

(TMX-71717) GEOMETRY

Blade row Stator Rotor

Pitch, cm 2.096 0.5621

Mean chord, cm 2.311 1.043

Height, cm 1.118 1.105

L.E. radius, cm 0.0635 0.0331

T.E. radius, cm 0.0191 0.0191

Number of blades 15 56

TABLE 3.—OPERATING CONDITIONS FOR ORIGINAL NASA DESIGNED TURBINE AND SCALED TURBINES

Parameters NASA 1979 design

Automotive turbine TMX-71717

Scaled NASA turbine

Inlet total temperature 1325 (°K) 1900 (°K)

Inlet total pressure 3.92 (atm) 17 (atm)

Mass flow rate 0.598 (kg/s) 2.166 (kg/s)

Rotor speed 6126.9 (rad/s) 7335.9 (rad/s)

Shaft power 22.615 [kW] 134.747 [kW]

Reaction 0.8013 0.8335

Total-to-total Polytrophic efficiency, percent 78.5385 87.6644

Referring to Table 3, the scaled turbine produces about 6 times the shaft power at 3.6 times the mass 

flow rate because of the 4.33 times higher intake total pressure. The other hand, the inlet flow stagnation 

temperature for the scaled turbine is only 1.433 times the NASA original turbine Figures 18 and 19 

present computed Mach number and static pressure carpet plots at 50 percent blade span for the NASA 

1975 designed and scaled turbines in the reference frame of each blade row. The contours indicate 

slightly higher Mach numbers in the scaled turbine compared to the original turbine. 

Figures 20 presents sample trajectories with color contouring of their velocity magnitude for 26 m 

alumina particles in the NASA designed turbine (Fig. 20(a)) and scaled NASA turbine (Fig. 20(b)) 

corresponding to particle seeding within 5 percent of the span near the hub and shroud. Most particles are 

seen to go through the stator blade passages with negligible interactions with the blade surfaces. However 

as they enter the rotor blade passages, they impact the rotor blades suction surface towards the leading 

edge in the NASA designed original turbine. The particles rebound with high circumferential velocities 

after their rotor blade suction surface impacts, which immediately drive them radially outward and back 

into the stator blade passages. In contrast, particles encounter multiple impact within the rotor blades row 

in the scaled NASA turbine but do not rebound back into the stator blade passages. 

Figures 21 and 22 present the predicted thermal barrier coating erosion rates for the NASA designed 

original turbine stator and rotor blade surfaces corresponding to uniform particle seeding (Fig. 21) and 

particle seeding within 5 percent of the span near the hub and shroud to represent suspended particles in 

the cooling secondary flow (Fig. 22). In both cases, high erosion rates are seen on rotor blades suction 

surface at the shroud close to the leading edge and on the stator blades suction surface at the casing near

trailing edge. The latter are associated with the repeated impacts by the particles that get trapped and 

bounce back and forth between the rotor and stator blade rows. Since the particles acquire the highest 

absolute velocities when they rebound from the rotor suction surface impacts, they produce the highest 

erosion rates by their subsequent impacts the stator blade suction surfaces. 
NASA/TM—2013-217857 16

(a) Original NASA Turbine

(b) Scaled NASA Turbine

Figure 20.—Sample three-dimensional particle 

trajectories ingested at 50 percent gas velocity 

within 5 percent of stator blade span near the hub 

and shroud ingested 50 percent gas velocity. 
NASA/TM—2013-217857 17

Figure 21.—TBC erosion rate predictions on blade 

surfaces in the NASA designed original turbine for 

particles ingested at 50 percent of the gas velocity

for uniform radial loading. 

 Figure 22.—TBC erosion rate prediction on blade 

surfaces in the NASA designed original turbine for 

particles ingested at 50 percent of the gas velocity

within 5 percent of stator blade span near the hub 

and shroud. 
NASA/TM—2013-217857 18

Lower erosion rates are predicted in the case of scaled turbine. Figures 23 and 24 present the 

computed thermal barrier coating erosion rates on the scaled NASA turbine stator and rotor blade surfaces 

for uniform particle seeding (Fig. 23) and for particle seeding within 5 percent of the span near the hub 

and shroud (Fig. 24) compared to the NASA designed original turbine (Fig. 22).

Table 4 gives a summary of the maximum local and overall predicted TBC erosion rates in both the 

NASA designed original and scaled turbine. The results in Table 4 indicate that both overall and 

maximum local TBC erosion rates are lower in the scaled turbine than the original turbine for the same 

particle loading distribution. This is attributed to the fact that very few particles bounce back from the 

rotor suction surface to impact the stator blades in the scaled NASA turbine. 

Figure 23.—TBC erosion rates on blade surfaces in the scaled 

NASA turbine for uniform particle loading at 50 percent of the 

gas velocity. 

 Figure 24.—TBC erosion rates on blade surfaces in the scaled 

NASA turbine for particles seeding within 5 percent of stator 

blade span near the hub and shroud at 50 percent of the gas 

velocity. 
NASA/TM—2013-217857 19

TABLE 4.—COMPUTED OVERALL BLADES AND MAXIMUM LOCAL TBC EROSION RATES 

ON THE COMMERCIAL TURBINE AND THE NASA DESIGNED TURBINE 

Turbine Particle radial 

loading% span

Maximum local erosion rate,

mg/g/cm2

Overall blade erosion rate,

mg/g

Stator Rotor

NASA original turbine 5% 60.21 (Stator suction side) 1.50 0.88

100% 20.43 (Stator suction side) 0.89 0.73

NASA scaled turbine 5% 21.36 (Rotor suction side) 0.28 0.43

100% 7.66 (Rotor suction side) 0.33 0.37

After we considered all the factors that can cause such a difference, we concluded that the high flow 

pressures in the case of scaled NASA turbines lead to lower values of the Reynolds number for the 

particle slip velocities relative to the gas. Subsequently lower drag forces by the gas flow on the particles 

through the stator leads to lower particle velocities at the stator exit in the case of the original NASA 

designed turbine. Referring to Figure 14, that leads to high particle incidence angles relative to the rotor 

suction surface and hence lower particle rebound velocities as indicated by the experimental measured 

restitution blade row coefficients presented in Figure 10. Subsequently these particles do not reenter the 

stator blade row nor impact the stator blade surfaces. 

It is interesting to compare results from the commercial APU turbine with uniform particle ingestion 

to those the scaled NASA designed automotive turbine with uniform particle ingestion. Both studies 

predict even distribution of erosion rate across the stator pressure surface with a significantly higher 

erosion rate on the rotor that is concentrated near the shroud. However particle trajectories in the NASA 

scaled turbine tended not to rebound from the rotor blade back into the stator blade row as in the original 

NASA and commercial APU turbine. This complex phenomenon is influenced by the several factors 

including the turbine blade airfoils, solidity, stator to rotor spacing as well as operating conditions. 

Summary

Experimental results are presented for particle surface interactions in the form of restitution 

coefficients after impacts with thermal barrier coated samples over the range of impact angles and erosion 

rates for flow velocities and temperatures ranging between 122 m/s (400 ft/s) to 366 m/s (1200 ft/s) and 

1871 °C (1600 °F) to 1093 °C (2000 °F), respectively. Predictions of TBC blade erosion by 26 m 

alumina particles are presented for the different single stage axial flow turbines namely a commercial 

APU, NASA designed turbine and for a scaled version of the NASA designed turbine for modern 

rotorcraft operating conditions. 

The highest overall thermal barrier coating erosion rates are always predicted on the rotor blade 

suction surface. On the other hand the pressure surfaces of the rotor blades always suffer the lowest 

thermal barrier coating erosion rates. The erosion rate of the thermal barrier coating on the stator suction 

surface caused by the impacts of particles that rebound from the rotor and reenter the stator blade passage 

is influenced by the restitution from the rotor blades suction. An important observation concerns the 

influence turbine flow pressure on the level of drag force on the suspended particles. This is important 

since the main aerodynamic force on the particles is associated with the drag coefficient that decreases 

with increased Reynolds number based on their slip velocity relative to gas flow.

References

1. Sirs, R.C., 1994, “The Operation of Gas Turbine Engines in Hot and Sandy Conditions-Royal Air 

Force Experiences in the Gulf War,” AGARD-CP-558, Paper No. 2. 

2. Mitchell, H.J. and Gilmore, F.R., 1982, “Dust-Cloud Effects on Aircraft Engines: Emerging Issues 

and New Damage Mechanisms,” RDA-TR-120012-001. 

3. Hamed, A., Tabakoff, W. and Wenglarz, R., 2006, “Erosion and Deposition in Turbomachinery,” 

Journal of Propulsion and Power 22(2), pp. 350–360. 
NASA/TM—2013-217857 20

4. Mund, M.G. and Guhna, H., 1970, “Gas Turbine Dust Air Cleaners,” ASME Paper 70-GT-104. 

5. Mann, D.L. and Wares, G.D., 1994, “Future Direction in Helicopter Engine Protection System 

Configuration,” AGARD-CP-588 Paper No. 4. 

6. McCoy, L., 1973, “The Coal Burning Gas Turbine Project, Report of the Interdepartmental Gas 

Turbine Steering Committee,” Australian Government Publishing Service. 

7. Hamed, A., Tabakoff, W. and Wenglarz, R., 1980, “Particulate Flow and Blade Erosion,” Von 

Karman Institute for Fluid Dynamics Lecture Series 1988–08. 

8. Meier, S.M. and Gupta, D.K., 1994, “The Evolution of Thermal Barrier Coatings in Gas Turbines 

Engine Applications,” Journal of Engineering for Gas Turbines and Power, 116, pp. 250–257. 

9. Miller, R.A., 1989, “Life Modeling of Thermal Barrier Coatings for Aircraft Gas Turbine Engines, in: 

D.E. Sokolowski (Ed.), Towards Improved Durability in Advanced Aircraft Engine Hot Sections,” 

NASA TM4087. 

10. Meier, S.M. et al., 1991, “Thermal Barrier Coating Life Prediction Model Development,” Phase II 

Final Report, Contract NAS3-23944, NASA CR 189111, (1991). 

11. Bose, S. and Masi-Marsin, J., 1995, “Thermal Barrier Coating Experience in Gas Turbine Engines at 

Pratt and Whitney,” Thermal Barrier Coating Workshop, NASA, pp. 63. 

12. Tabakoff, W. and Wakeman, T., 1979, “Test Facility for Material Erosion at High Temperature,” 

ASTM Special Publication, 664, pp. 123–135. 

13. Tabakoff, W., 1989, “Investigation of Coating at High Temperature for Use in Turbomachinery,” 

Surface and Coating Technology, 940, pp. 97–115. 

14. Nicolls, J.R., Deakin, M.J. and Rickerby, D.S., 1999, “A Comparison Between the Erosion Behavior 

of Thermal Spray and Electron Beam Physical Vapor Deposition Thermal Barrier Coatings,” Wear, 

233-235, pp. 352–361. 

15. Hussein, M.F. and Tabakoff, W., 1973, “Dynamic Behavior of Solid Particles Suspended by Polluted 

Air Flow in a Turbine Stage,” Journal of Aircraft 10(7), pp. 434–440. 

16. Hamed, A., Tabakoff, W., Rivir, R.B., Das K. and Arora, P., 2005, “Turbine Blade Surface 

Deterioration by Erosion,” Journal of Turbomachinery, 127, pp. 445–452. 

17. Hamed, A. and Tabakoff, W., 1994, “Experimental and Numerical Simulations of the Effects of 

Ingested Particles in Gas Turbine Engines,” AGARD-CP-558, Erosion, Corrosion and Foreign Object 

Effects in Gas Turbines. 

18. Roelke, R.J., and McLallin, K.L., “The Aerodynamic Design of Compressor-Drive Turbine for use in 

a 75 kW Automotive Engine,” NASA Technical Memorandum. TMX-71717, 1975. 

19. Swar, R., 2009, “Particle Erosion of Gas Turbine Thermal Barrier Coating,” M.S. Thesis, School of 

Aerospace Systems, University of Cincinnati, Cincinnati, OH. 

20. Tabakoff, W., Hamed, A. and Beacher, B., 1983, “Investigation of Gas Particle Flow in an Erosion 

Wind Tunnel,” Journal of WEAR, Vol. 86, Apr, pp. 73–88. 

21. Gunaraj, J.A., Tabakoff, W., and Shastri, S., 1997, “Experimental Investigation of Solid Particle 

Interaction in Particulate Flow,” ASME Fluids Engineering Division Summer Meeting Paper No. 

FEDSM97-3628. 

22. Gunaraj, J.A., 1997, “Study of the Effect Particle-Particle Interactions in the Particulate Flow,” PhD. 

Thesis, University of Cincinnati. 

23. Shastri, S., 1999, “Particle Rebound Characteristics of Turbomachinery Cascade Leading Edge 

Geometry,” PhD. Thesis, University of Cincinnati. 

24. ANSYS CFX, Release 12.1, User Manual, 2009. 

25. Pointwise, Release 16.04, User Manual, 2011. 

26. Woggon, N., “Particle Erosion of a Turbine With Restitution Analysis,” M.S. Thesis, University of 

Cincinnati, 2012. 

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Combined Experimental and Numerical Simulations of Thermal Barrier Coated Turbine 

Blades Erosion 

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Hamed, Awatef; Tabakoff, Widen; Swar, Rohan; Shin, Dongyun; Woggon, Nathanial; Miller, 

Robert 

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14. ABSTRACT

A combined experimental and computational study was conducted to investigate the erosion of thermal barrier coated (TBC) blade surfaces 

by alumina particles ingestion in a single stage turbine. In the experimental investigation, tests of particle surface interactions were 

performed in specially designed tunnels to determine the erosion rates and particle restitution characteristics under different impact 

conditions. The experimental results show that the erosion rates increase with increased impingement angle, impact velocity and

temperature. In the computational simulations, an Euler-Lagrangian two stage approach is used in obtaining numerical solutions to the threedimensional compressible Reynolds Averaged Navier-Stokes equations and the particles equations of motion in each blade passage 

reference frame. User defined functions (UDF) were developed to represent experimentally-based correlations for particle surface 

interaction models which were employed in the three-dimensional particle trajectory simulations to determine the particle rebound 

characteristics after each surface impact. The experimentally based erosion UDF model was used to predict the TBC erosion rates on the 

turbine blade surfaces based on the computed statistical data of the particles impact locations, velocities and angles relative to the blade 

surface. Computational results are presented for the predicted TBC blade erosion in a single stage commercial APU turbine, for a NASA 

designed automotive turbine, and for the NASA turbine scaled for modern rotorcraft operating conditions. The erosion patterns in the 

turbines are discussed for uniform particle ingestion and for particle ingestion concentrated in the inner and outer 5 percent of the stator 

blade span representing the flow cooling the combustor liner. 

15. SUBJECT TERMS

Coatings; Erosion; Computational fluid dynamics 

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