Harold E. Addy, Jr.

Glenn Research Center, Cleveland, Ohio

Andy P. Broeren

University of Illinois at Urbana-Champaign, Urbana, Illinois

Joseph G. Zoeckler

Glenn Research Center, Cleveland, Ohio

Sam Lee

QSS Group, Inc., Cleveland, Ohio

A Wind Tunnel Study of Icing Effects

on a Business Jet Airfoil

NASA/TM—2003-212124

February 2003

AIAA–2003–0727
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Harold E. Addy, Jr.

Glenn Research Center, Cleveland, Ohio

Andy P. Broeren

University of Illinois at Urbana-Champaign, Urbana, Illinois

Joseph G. Zoeckler

Glenn Research Center, Cleveland, Ohio

Sam Lee

QSS Group, Inc., Cleveland, Ohio

A Wind Tunnel Study of Icing Effects

on a Business Jet Airfoil

NASA/TM—2003-212124

February 2003

National Aeronautics and

Space Administration

Glenn Research Center

Prepared for the

41st Aerospace Sciences Meeting and Exhibit

sponsored by the American Institute of Aeronautics and Astronautics

Reno, Nevada, January 6–9, 2003

AIAA–2003–0727
Available from

NASA Center for Aerospace Information

7121 Standard Drive

Hanover, MD 21076

National Technical Information Service

5285 Port Royal Road

Springfield, VA 22100

Available electronically at http://gltrs.grc.nasa.gov

The Propulsion and Power Program at

NASA Glenn Research Center sponsored this work.

Acknowledgments

The authors wish to express their sincere gratitude to Mr. David Justavick and the technicians at the

NASA Glenn Icing Research Tunnel, model craftsmen Mr. Daniel Gura and Mr. Robert Reminder,

Ms. Pamela Phillips, Mr. William Sewall and the technicians at the NASA Langley Low Turbulence

Pressure Tunnel without whose outstanding support, expertise, and dedication this work

could not have been completed.
NASA/TMó2003-212124 1 

A WIND TUNNEL STUDY OF ICING EFFECTS ON A BUSINESS JET AIRFOIL 

Harold E. Addy, Jr.,∗

 Andy P. Broeren,Ü Joseph G. Zoeckler,á and Sam Lee§

∗á

National Aeronautics and Space Administration, Glenn Research Center, Cleveland, Ohio Ü

University of Illinois at Urbana-Champaign, Urbana, Illinois §

QSS Group, Inc., Cleveland, Ohio 

ABSTRACT

 Aerodynamic wind tunnel tests were 

conducted to study the effects of various ice accretions 

on the aerodynamic performance of a 36-inch chord, 

two-dimensional business jet airfoil. Eight different ice 

shape configurations were tested. Four were castings 

made from molds of ice shapes accreted in an icing 

wind tunnel. Two were made using computationally 

smoothed tracings of two of the ice shapes accreted in 

the icing tunnel. These smoothed profiles were then 

extended in the spanwise direction to form a twodimensional ice shape. The final two configurations 

were formed by applying grit to the smoothed ice 

shapes. The ice shapes resulted in as much as 48% 

reduction in maximum lift coefficient from that of the 

clean airfoil. Large increases in drag and changes in 

pitching moment were also observed. The castings and 

their corresponding smoothed counterparts yielded 

similar results. Little change in performance was 

observed with the addition of grit to the smoothed ice 

shapes. Changes in the Reynolds number (from 3×106

to 10.5×106

) and Mach number (from 0.12 to 0.28) did 

not significantly affect the iced-airfoil performance 

coefficients. 

NOMENCLATURE

α Airfoil angle-of-attack 

c Airfoil chord length 

Cd Drag coefficient 

Cl

 Lift coefficient 

Cl,ma Maximum lift coefficient, coincident with αstall 

Cm Quarter-chord pitching-moment coefficient 

Ma Freestream Mach number 

Re Reynolds number based on chord 

x Chordwise position along airfoil 

LWC Liquid water content 

MVD Median volumetric diameter 

IRT Icing Research Tunnel 

LTPT Low Turbulence Pressure Tunnel 

INTRODUCTION

 In the 1990’s, the Icing Branch at NASA’s 

Glenn Research Center, in conjunction with the FAA’s 

Technical Center, embarked on an effort to investigate 

the range of shapes and sizes of inflight ice that might 

accrete on various newer airfoils. At that time, the 

existing ice shape information was limited primarily to 

older airfoil sections or simpler airfoils such as the 

symmetric NACA-0012. Although many of these 

airfoils are still in use, there was a need to find whether 

or not the newer airfoil designs accreted ice and 

affected their performance differently. Among the 

goals of the effort were to collect ice shape data for 

computer ice accretion code validation and to study the 

effects of the ice on aerodynamic performance. 

 Industry was surveyed to find airfoils suitable 

for the study. Three airfoils were selected: one 

representative of a commercial transport horizontal 

tailplane, one representative of a business jet main 

wing, and one representative of a general aviation 

aircraft main wing. A 36-inch chord, two-dimensional 

model of each airfoil was built and tested under various, 

FAR Part 25, Appendix C, icing conditions in NASA 

Glenn’s Icing Research Tunnel (IRT). The ice shapes 

were recorded using photographs, ice tracings, and, for 

selected shapes, molds. Aerodynamic performance 

measurements were also made for a portion of the test 

runs. Results of these tests were documented by Addy.1

 

Wright and Rutkowski2

 conducted an extensive study of 

how these ice shapes, along with others, compared with 

those predicted by LEWICE 2.0,3

 a computer ice 

accretion code. In general, agreement between the ice 

tracings and the corresponding computationally 

predicted shapes was quite good. 

 As mentioned earlier, some aerodynamic 

performance measurements were made in the IRT. 

However, the icing cloud in the IRT did not always fill 

the entire test section, resulting in ice shapes that 

tapered off near the walls. These non-representative ice 

shapes near the walls might have significantly altered 

the performance measurements. Therefore, molds were 

 

∗

 Research Engineer, Icing Branch, Member AIAA. 

Ü

 Post-Doctoral Research Scientist, Department of Aeronautical and Astronautical Engineering, Member AIAA. 

á

 Operations Engineer, Space Mechanical Engineering Branch. 

§

 Research Engineer, Icing Branch, Member AIAA. 
NASA/TMó2003-212124 2 

made of the ice at the center of the airfoil for several 

selected ice shapes. Castings were made from these 

molds and placed on a similar model in an aerodynamic 

wind tunnel better suited to make the aerodynamic 

performance measurements. 

 Aerodynamic performance measurements 

were made with various ice shapes on both the general 

aviation model and the business jet model. Results 

from the tests with the general aviation model were 

documented by Addy1

 and by Addy and Chung.4 This 

paper presents results from the aerodynamic 

performance measurements made with the business jet 

model. Aerodynamic performance measurements have 

not been conducted in an aerodynamic tunnel with the 

commercial transport model. The performance 

measurements made in the aerodynamic wind tunnel 

are more comprehensive and are more accurate than 

those made in the IRT. Therefore, the results presented 

here are from the aerodynamic wind tunnel only. 

TEST DESCRIPTION

FACILITY

 The wind tunnel used for these tests was NASA 

Langley Research Center’s Low Turbulence Pressure 

Tunnel (LTPT). Shown schematically in Fig. 1, the 

LTPT is a closed return wind tunnel that is principally 

used for two-dimensional airfoil testing.5,6 It can be 

operated at stagnation pressures from near vacuum to 

147 psia and over a Mach number range of 0.05 to 0.40. 

The maximum Reynolds number is dependent upon 

Mach number. For example, the maximum Reynolds 

number per foot is 15×106

 at a Mach number of 0.22. A 

heat exchanger and nine turbulence reduction screens 

are located in the settling chamber. The contraction 

ratio is 17.6:1 and the test section dimensions are 36-

inches wide by 90-inches high by 90-inches long. The 

tunnel was designed for two-dimensional airfoil testing 

with model chord lengths up to 36-inches.5

 The free 

stream turbulence levels have been measured at less 

than 0.1% for all operating conditions.6

 

The airfoil model was supported horizontally 

across the width of the test section between two 40-inch 

diameter circular endplates. The endplates were flush 

with the sidewalls and rotated for angle-of-attack 

adjustment. They also contained a section of porous 

plate for sidewall boundary layer control. This sidewall 

venting system was originally developed for testing 

high lift airfoil configurations. A detailed description 

of the system is given by Paschal, et al.7

 During this 

study, some runs were performed with and without 

sidewall venting. There was very little difference in the 

results. 

 The LTPT was equipped with a threecomponent force balance. However, it was designed 

for operation with high lift systems and for higher 

dynamic pressures than were run in this experiment. 

Therefore, the data from the force balance were deemed 

unreliable except at the higher dynamic pressures, and 

the lift and pitching moment data were obtained from 

the integration of surface static pressures. These data 

were compared to the force balance data for large 

values of dynamic pressure and excellent agreement 

was observed for both clean and iced configurations. 

Therefore, no force balance data are presented here. 

The pitching moment coefficient reported here was 

determined for the quarter chord location. Drag 

coefficients were calculated from wake pressures 

measured with a wake probe using the standard 

momentum deficit method. Due to the rather long 

length of time required to complete a wake traverse, the 

number of angles-of-attack where wake surveys were 

conducted was limited. Corrections to the integrated 

performance coefficients accounting for solid and wake 

blockage and streamline curvature were applied to the 

data during post processing using the methods of Allen 

and Vincenti.8

AIRFOIL

 The business jet main wing airfoil used in 

these tests was based upon a GLC-3059

 airfoil. This 

airfoil was designed for low transonic drag and had a 

maximum thickness to chord ratio of 8.7%. Its cross 

section is shown in Fig. 2. The two-dimensional model 

used in these tests had a 36-inch chord and 36-inch 

span. It was machined from solid aluminum. Channels 

were made for pressure taps and tubing. The pressure 

taps were installed in both the chordwise and spanwise 

directions. The model was also designed with 

removable leading edges. There were two leading edge 

designs: a baseline leading edge for the clean model 

testing and an alternate leading edge that allowed 

installation of the castings that replicated the ice 

accreted in the IRT. The removable leading edge 

design resulted in a spanwise-running seam on the 

upper surface at x/c = 0.028 and on the lower surface at 

x/c = 0.215. Another removable section was required 

to facilitate model installation and this resulted in an 

additional upper surface seam at x/c = 0.111. 

Alignment of these edges was very good and no 

obvious detrimental effects in the performance data 

were observed. 

ICE SHAPES

 Ice shapes were accreted on the GLC-305 airfoil in 

the IRT over a range of icing conditions selected from 

the FAA’s FAR Part 25-Appendix C Atmospheric Icing 

Conditions. Even though the 36-inch chord of the 

model was not typical of a business jet main wing, the 

icing conditions were not scaled. At the time of the ice 

accretion tests, methods of icing scaling for model size 
NASA/TMó2003-212124 3 

had not been validated. Airspeeds and model attitudes 

typical of a business jet aircraft were used. Molds10

were made of selected ice shapes formed during these 

tests. Castings were made from four of these molds for 

the aerodynamic tests in the LTPT. Figure 3 shows 

tracings made at the model centerline of the four ice 

shapes. The molds were made at center-span of the 

model and covered from ten to twelve inches of span of 

the ice shape. This mold covered enough of the span of 

the ice to sufficiently reproduce the variation in ice 

along the span of the model. The spanwise variation in 

ice shape was small such that when three to four 

castings were made from the same mold and placed 

side by side across the span of the LTPT model, no 

significant discontinuities were present. The icing 

conditions under which these ice shapes were accreted 

are given in Table 1. 

Table 1. Icing Conditions 

 Exposure 

Ice Vel AOA Tt MVD LWC Time 

Shape kts. deg deg F m-6 g/m3 min. 

202 175 6 30 20 0.54 2.0 

212 175 6 12 20 0.40 16.7 

944 175 4 30 20 0.54 22.5 

904 175 6 30 20 0.54 22.5 

In addition to the castings, two smooth ice 

shapes of constant cross section were manufactured and 

tested. These smoothed shapes were made from 

tracings of ice that was also used to make the castings. 

Separate test runs were made using the same icing 

conditions. One was used to make the ice tracing and 

the other was used to make the ice mold. The tracings 

were first digitized using a digitizing tablet, and then 

smoothed using a computational routine known as 

SmaggIce.11 The resulting smoothed tracings are shown 

in Fig. 4. These smoothed tracings were then projected 

in the spanwise direction with a computer aided design 

program and then manufactured using a laser sintering 

rapid prototyping technique. The effects of the 

smoothing process are not apparent in the cross sections 

shown in Fig. 4 for several reasons. First, the tracings, 

made with a pencil, did not pick up the finer features 

(those less than the diameter of the pencil lead in 

dimension) of the ice. Secondly, the amount of 

smoothing applied by SmaggIce was very slight, only 

removing the sharpest corners. Finally, the smoothing 

in the spanwise direction, accomplished by extending 

the smoothed cross section laterally, was significant and 

constituted a major difference between the a casting and 

corresponding smoothed ice shape. This feature does 

not appear in the cross section views of the ice shapes. 

 

The smooth ice shapes (212 and 944) were 

tested in both the as-manufactured condition and with 

No. 30 grit applied to their surfaces. This exercise was 

conducted as a cursory look at whether aerodynamic 

performance could be significantly altered by the 

presence of a smaller, ice-like feature on the smooth ice 

shape. No. 30 grit was chosen because its size is 

similar to that of typical real ice roughness. An 

adhesive was brushed onto the smooth ice shapes and 

grit was applied in a sparse pattern similar to real ice. 

 The ice shape castings were mounted on the 

alternate leading edges. Since the castings were made 

from molds of the ice at the center of the model in the 

IRT, the castings were repeated such that the ice shapes 

spanned the entire width of the model. In addition, an 

“instrumentation slice” was installed near the model 

mid-span. The instrumentation slice was cut out of 

stainless steel to match the smoothed ice shape crosssection and had pressure taps distributed around the ice 

shape contour. This allowed for a good approximate 

measurement of the pressure distribution around the ice 

shape and provided pressures for determination of the 

lift and pitching moment coefficients. The baseline 

model had 93 static pressure orifices along the main 

chordwise row and 20 orifices in a spanwise row 

located at x/c = 0.70 on the upper surface. The 

photograph in Fig. 5 shows the model mounted in the 

test section, with ice shape and pressure instrumentation 

slice attached to the airfoil leading edge. 

AERODYNAMIC CONDITIONS

 The test matrix for the LTPT tests was selected 

to yield a broad range of Reynolds and Mach numbers 

with the high end being applicable to those experienced 

by a modern business jet flying in typical icing 

conditions as constrained by the limitations of the 

facility. Reynolds numbers ranged from 3×106

 to 

10.5×106

, while Mach numbers ranged from 0.12 to 

0.28. The test matrix was designed to isolate the effects 

of Reynolds and Mach number. Therefore, there is a 

Reynolds number variation from 3.0×106

 to 10.5×106 at 

a fixed Mach number of 0.12. Similarly, there is a 

Mach number variation from 0.12 to 0.28 at constant 

Reynolds numbers of 6.0x106

 and 10.5×106. The 

combinations of Mach and Reynolds numbers are given 

in Table 2. Note that not all combinations of Mach and 

Reynolds numbers were run for all ice shape 

configurations due to tunnel and test time limitations. 

 The below atmospheric pressure capability of 

the LTPT was not utilized for this test. Sidewall 

venting was typically employed when testing at above 

atmospheric pressures in the tunnel. 
NASA/TMó2003-212124 4 

Table 2. LTPT Airflow Conditions 

Mach No. 

Re No. 0.12 0.21 0.28 

3x106 X 

3.5 X 

4.5 X 

6 X X X 

7.5 X X 

10.5 X X X 

RESULTS AND DISCUSSION

CLEAN AIRFOIL RESULTS

The clean performance of the GLC-305 airfoil 

generally exhibited classic Reynolds and Mach number 

behavior. Establishing the clean baseline performance 

of this airfoil was important since no experimental data 

of this kind were available in the public domain to the 

authors’ knowledge. The effect of Reynolds number on 

the performance at a constant Mach number of 0.12 is 

summarized in Fig. 6. There was a significant increase 

in lift curve slope, maximum lift and stalling angle-ofattack between Re = 3.0×106

 and 6.0×106. The Cl,max

increased from 1.09 to 1.35 over this Re range. As the 

Reynolds number was increased to 10.5×106

, the stall 

type changed from a leading edge type to a trailing edge 

stall. Using the stall type definitions established by 

McCullough and Gault,12 the sharp drop off in Cl

 for Re

= 3.0×106

 and 6.0×106 post maximum lift was indicative 

of a leading edge stall caused by the bursting of small 

laminar separation bubble that formed near the airfoil 

leading edge. For Re = 10.5×106

 the stall characteristics 

were consistent with the trailing edge type caused by 

the gradual forward movement of boundary layer 

separation from the trailing edge with increasing angleof-attack. The lift and pitching moment data indicated 

that the GLC-305 airfoil was a front loaded section that 

was not highly cambered. The zero lift angle-of-attack 

was only about –1 deg. and the pitching moment 

coefficients were small and nearly constant over the 

linear range of the lift curve. 

The variation in the section drag coefficients 

over this Reynolds number range is also summarized in 

Fig. 6. An apparent anomaly exists in these data since 

the drag values for α = 0 and 2 deg. were lowest at 

Re = 3.0×106

. For angles-of-attack greater than two 

deg., the Cd at this Reynolds number was higher than 

the Cd for Re = 6.0×106

 and 10.5×106, as expected. The 

apparent anomaly in the Re = 3.0×106

 case is not 

unprecedented. For example, Abbot and von 

Doenhoff13 show drag data for the NACA 0006 and 

0009 airfoils that exhibited analogous behavior between 

Re = 3.0×106

 and 6.0×106. The NACA airfoils had 

thickness ratios of 6% and 9%, respectively, while the 

GLC-305 airfoil had a thickness to chord ratio of 8.7%. 

This behavior in the drag was also investigated using 

XFOIL14 and similar results were observed. Longer 

runs of laminar flow along the airfoil surface are 

expected at the lower free stream Reynolds number, 

provided the flow does not separate and form a bubble. 

Usually, the reduced drag benefit of an extended 

laminar boundary layer is offset by separation of the 

turbulent boundary layer near the trailing edge. 

However, for the low angles-of-attack on these thin 

airfoils, the boundary layer separation was minimal 

such that the drag was lower for Re = 3.0×106

. 

Compressibility effects on the performance of 

the clean GLC-305 airfoil were investigated over a 

Mach number range of 0.12 to 0.28 at a constant 

Re = 10.5×106

 and these data are shown in Fig. 7. As 

with the Reynolds number behavior, classic Mach 

number trends were also observed. Most noticeably, 

there was a significant reduction in maximum lift 

coefficient and stalling angle-of-attack between 

Ma = 0.21 and 0.28. Unfortunately, data at Ma = 0.28 

could not be acquired for angles-of-attack larger than 

12 deg. This was a limitation imposed by the facility 

due to the unsteady flow associated with the stall and 

the large dynamic pressure associated with this Mach 

number. The classic increase in lift curve slope with 

increasing Mach number was not as apparent as was the 

analogous effect on the pitching moment. The effect of 

increasing Mach number on the drag was most apparent 

for angles-of-attack larger than six deg., but still 

followed classic trends. 

These independent effects of Reynolds and 

Mach number on the clean airfoil performance were 

consistent with previous LTPT experiments. For 

example, Ladson15 analyzed data from the testing of a 

NACA 0012 airfoil having a 24-inch chord, acquired 

with solid sidewalls (no sidewall boundary layer 

control) over a large range of Reynolds and Mach 

numbers. The similarity in the variation in maximum 

lift with Reynolds number at constant Mach number is 

shown in Fig. 8. These similarities exist despite the 

differences in the airfoil section and model geometries. 

The variation with Mach number was nearly identical, 

with each airfoil’s Cl,max decreasing by about 0.30 over 

the range indicated in Fig. 9. The large decrease in 

maximum lift between Ma = 0.21 and 0.28 

corresponded to an increase in the minimum pressure 

(Cp,min) measured on the airfoil. The “critical Cp” shown 

in Fig. 10 is the Cp corresponding to a local Mach 

number equal to 1.0 for a given free stream Mach 

number.16 The minimum Cp for the Ma = 0.28 case in 

Fig. 10, was at the stalling angle-of-attack for this case. 

The magnitude reduction of the minimum pressure peak 

for local sonic Mach numbers is consistent with 

previous studies as the airfoil is unable to support the 

transonic flow.17,18,19 
NASA/TMó2003-212124 5 

ICED AIRFOIL RESULTS

 The effects of the different cast ice shapes on 

lift coefficient are shown in Fig. 11a. It is interesting 

to note that the two-minute glaze ice shape 

(202 casting) has nearly the same effect on lift 

coefficient as the 16.7-minute rime shape (212 casting). 

In both cases, the maximum lift coefficient is reduced 

by approximately 20%. The two 22.5-minute glaze ice 

shapes (904 and 944 castings) also have very similar 

effects on lift coefficient even though the two shapes 

were formed at different angles-of-attack, but under the 

same icing conditions. Here the reduction in maximum 

lift coefficient is much greater, approximately 48%. 

 The effects of the cast ice shapes on drag and 

pitching moment are shown in Figs. 11b and 11c. Here 

again, the similarities between the two-minute glaze 

shape and the 16.7-minute rime shape and the two 

22.5-minute glaze shapes are apparent, although the 

difference in pitching moment between the two-minute 

glaze shape and 16.7-minute rime shape is more 

evident. Because it acts as a leading edge extension, 

the rime shape generates more lift before stall than the 

two-minute glaze shape. This effect can be seen in the 

pitching moment coefficient as the rime ice shape 

produces a more nose up moment condition. 

 The effects of varying Reynolds number on 

three of the cast ice shapes are shown in Figs. 12, 13, 

and 14. In contrast to the clean ice shape where a 

strong dependence of performance on Reynolds number 

was observed, little or no dependence of performance 

on Reynolds number was observed for these ice shapes. 

Only on the smallest ice shape, the two-minute glaze 

shape, was a very small difference in maximum lift 

coefficient observed. For a clean airfoil, maximum lift 

and stall characteristics are largely determined by 

boundary layer transition and separation that are in turn 

governed by Reynolds number. For an iced airfoil, 

boundary layer separation is generally fixed by the size 

and location of the ice shape. Therefore, it follows that 

iced airfoil performance characteristics are less 

sensitive to changes in Reynolds number. 

 Mach number effects on the same three cast 

ice shapes are shown in Figs. 15, 16, and 17. Similar to 

the clean airfoil, small but noticeable Mach number 

effects are observed for the ice shapes. This behavior 

has also been observed in previous tests. While this 

type of compressibility effect has been widely observed 

at Mach numbers greater than 0.30, the effect is 

generally negligible at Mach numbers below this level. 

The present data show that small, but noticeable 

compressibility effects can be measured at lower Mach 

numbers as well. The more significant point here is 

that Mach number, a measure of airspeed relative to the 

speed of sound, tends to have a similar effect on the 

airfoil performance coefficients regardless of whether 

or not it is ice contaminated. This is in contrast to the 

Reynolds number, a measure of viscous effects, that has 

a large effect on the clean airfoil, but virtually no effect 

when it is ice contaminated. 

The independent effects of Reynolds and 

Mach numbers on the maximum lift performance of the 

iced GLC-305 airfoil are summarized in Figs. 18 and 

19. The Reynolds number trends given in Figs. 12 

through 14 were indicative of the other configurations 

tested as well. In particular, Fig. 18 shows that 

Reynolds number changes did not affect the difference 

in maximum lift between the airfoil with the 

944 casting and the airfoil with the 944 smoothed 

shape. On the other hand, the minor increase in Cl,max

from Re = 3.0×106

 to 10.5×106 for the airfoil with the 

212 casting, did not occur for the smoothed ice shape. 

Figure 19 shows that the maximum lift for all of the 

iced configurations tended to decrease as the Mach 

number increased and this trend is consistent with the 

clean case. The opposite is true of the Reynolds 

number effect where the iced configurations tended to 

have maximum lift values that were, for the most part, 

insensitive to Reynolds number changes. 

 The Reynolds and Mach number trends 

observed for the iced airfoil cases are consistent with 

previous research. For example, Morgan, et al.20 carried 

out iced airfoil performance measurements on a multielement super-critical airfoil. A large glaze ice 

simulation made of wood was tested as well as 

simulated frost that consisted of 70-grit roughness. 

Performance measurements were carried out with all 

airfoil elements nested (cruise configuration) at a 

constant Mach number of 0.20 with Re = 3.0×106

 to 

12.0×106

. The results showed minor changes in 

maximum lift over the entire Reynolds number range 

for the iced airfoil cases. In another study, Addy and 

Chung4

 tested glaze ice simulations on an NLF-0414 

airfoil in the LTPT. The simulations consisted of 

castings produced from ice accretions and 

corresponding smoothed shapes (similar to the present 

study). The ice accretions resulted from a 6-minute and 

22.5 minute exposure to the same cloud conditions. For 

three of the four iced airfoil configurations tested, there 

was virtually no variation in Cl,max for Re = 4.6×106

 to 

10.5×106

, at a constant Mach number of 0.21. The 

authors also reported similar Mach number trends in 

maximum lift behavior in the iced airfoil configurations 

at Re = 6.4×106

. More recently, Broeren, et al.21 carried 

out performance measurements on a NACA 23012 

airfoil with intercycle ice castings. In this case there 

was a slight increase in Cl,max (less than 0.05) between 

Re = 2.0×106

 and 3.5×106, but no further increase in 

Cl,max up to Re = 10.5×106

 was observed for all four iced 

airfoil configurations. Also similar to the present data, 

the iced airfoil maximum lift coefficient tended to 

decrease with increasing Mach number at a constant 

Re = 10.5×106

. 
NASA/TMó2003-212124 6 

 Some insight into the iced airfoil Reynolds and 

Mach number effects can be gleaned from the surface 

pressure distributions. Figure 20 compares the Cp

distributions on the airfoil with the 944 smoothed shape 

at 4 deg. angle-of-attack for two different Reynolds 

numbers. Both curves are indicative of a separation 

bubble flowfield aft of the large glaze ice shape (e.g., 

see Bragg, et al.22). The constant pressure region on the 

upper surface between x/c = -0.02 and 0.16 indicates 

flow separation from the ice shape. As the Cp increases, 

transition occurs in the separated shear layer and the 

bubble reattaches further downstream. The pressure 

recovery region for Re = 3.5×106

 is slightly upstream of 

the recovery region for Re = 10.5×106

. This may imply 

that the separation bubble reattachment location was 

also slightly upstream at the lower Reynolds number. 

Figure 21 shows the effect of increasing the Mach 

number at constant Reynolds number. In this case, the 

increase in Mach number also has a small effect on the 

Cp values in the constant pressure region. Comparison 

of the recovery region also implies that the bubble 

reattachment location may be further downstream for 

the higher Mach number case. 

 As was mentioned previously, for two of the 

ice shapes, the 22.5-minute glaze shape 944 and the 

16.7-minute rime shape 212, smoothed, twodimensional ice shapes were made. Performance 

results for the castings and their smoothed counterparts 

are given in Figs. 22 and 23. The differences between 

the lift, drag, and pitching moment coefficients for the 

cast and smoothed, 22.5-minute glaze ice 944 shapes 

are very small, as shown in Fig. 22. There is more of a 

difference in the performance curves between the cast 

and smoothed, 16.7-minute rime ice shapes 212 at the 

higher angles-of-attack. For instance, the maximum 

lift coefficient for the smoothed shape is 5% lower than 

that for the cast ice shape. Here, the three-dimensional 

features of the cast ice shape may have enhanced 

mixing of the shear layer, promoting reattachment that 

resulted in slightly higher lift. This difference is small, 

however, particularly in comparison to the difference 

between these ice shapes and the clean airfoil, where 

there was a 22% reduction in maximum lift coefficient. 

Whether this difference is small enough for modeling 

purposes is dependent upon the application. 

 These results for the cast versus smoothed ice 

shapes are, however, in contrast to results found with 

the general aviation airfoil as described in Reference 3. 

In that study, the maximum lift coefficients differed by 

9% between the cast and smoothed shapes for both a 

six-minute and a 22.5-minute glaze ice shape. Perhaps 

more significantly, the cast ice shapes generated greater 

performance degradation than the smoothed ice shapes, 

which was unexpected. Apparently, the shear layer 

mixing that promotes reattachment either did not occur 

or was overridden by another unexplained 

phenomenon. The two airfoils used in these two studies 

do have different designs and performance 

characteristics that could contribute to the different 

results. However, different degrees of smoothing were 

also employed. Using the smoothing technique of 

SmaggIce,10 50% control points were used for the 

six-minute ice shape in the general aviation airfoil study 

and 30% control points were used for the 22.5-minute 

glaze ice shape. For the smoothed ice shapes in this 

business jet airfoil study, 100% control points were 

used. (In SmaggIce, the higher the percentage of 

control points assigned to the digitized ice shape, the 

closer the smoothed shape is to the original, digitized 

ice tracing.) While further investigation is warranted, 

these results suggest that an overly smoothed ice shape 

can lead to less precise results. 

 To further investigate the effects of smaller ice 

features on aerodynamic performance, a sparse 

distribution of No. 30 grit was applied to each of the 

smoothed ice shapes while installed on the model in the 

tunnel and the set of aerodynamic conditions were 

rerun. Results of these tests are shown in Figs. 24 and 

25. In neither the 16.7-minute rime ice shape case nor 

the 22.5-minute glaze ice shape 944 case, did the 

addition of the grit have much of an effect, if any, on 

the performance results. There was a slight increase in 

drag at angles-of-attack of zero and below for the 

smoothed glaze ice shape 944. For the smoothed rime 

ice shape, the presence of the grit generated a difference 

in the pitching moment curves in the post stall region. 

Otherwise, the performance curves for the grit and nogrit cases lay on top of each other. It should be noted 

that the process of selecting and applying material in 

order to reproduce the effects of ice roughness is more 

an art than a science. It is possible that a different grit 

size and/or distribution could produce slightly different 

results in performance degradation. 

CONCLUDING REMARKS

 Wind tunnel tests were conducted with a twodimensional business jet airfoil. Four different cast ice 

shapes were attached to the leading edge of the airfoil 

to determine the degradation in lift, drag, and pitching 

moment coefficients due to various types of ice 

accretions on this model. In addition, two-dimensional, 

smoothed versions of two of the cast ice shapes 

(212 and 944) were also tested to help understand 

differences that might be caused by the three 

dimensional characteristics of natural ice accretions. 

Grit was also applied to the smoothed ice shapes and 

the tests repeated to investigate the effects of smaller 

ice features on aerodynamic performance. 

 Results of the tests showed that the presence of 

the ice caused significant performance degradation. 

A two-minute glaze ice accretion reduced the maximum 

lift coefficient some 22% from the clean wing 

performance at Ma = 0.21 and Re = 7.5x106

. 
NASA/TMó2003-212124 7 

Interestingly, a 16.7-minute rime accretion resulted in 

similar lift coefficient degradation and lift coefficient 

versus angle-of-attack curve. Two 22.5-minute glaze 

ice shapes, accreted at the same icing conditions, but at 

two different model attitudes, 4 and 6 degrees, resulted 

in a 48% reduction in maximum lift coefficient and 

nearly identical lift coefficient versus angle-of-attack 

curves, also at Ma = 0.21 and Re = 7.5x106

. 

 The tests were conducted in the Low 

Turbulence Pressure Tunnel at NASA Langley, which 

allowed the independent variation of Mach and 

Reynolds numbers. As expected, the clean airfoil 

exhibited varying performance with varying Re = 3.0 to 

10.5x106

 at constant Ma = 0.12. However, with the ice 

shapes attached to the leading edge of the model, there 

was no variation in performance as a result of changing 

Reynolds number over the same range. This result 

supports the notion that ice accretions dominate the 

boundary layer behavior on iced airfoils. Mach number 

effects (from 0.12 to 0.28 at constant Re = 10.5x106

), 

although rather small, were observed for both the clean 

airfoil and with all the ice shapes used in the tests. 

 The smoothed ice shapes 212 and 944 

produced performance results very similar to their cast 

ice shape counterparts in this study. This result was in 

contrast to previous test results with smoothed ice 

shapes on the general aviation model. In addition to the 

different airfoils used, a different amount of smoothing 

was applied to the ice shapes in the two studies. Less 

smoothing was applied to the ice shapes in this study 

with the business jet airfoil, resulting in smoothed ice 

shapes much closer in profile to their cast ice shape 

counterparts.

 Finally, No. 30 grit applied to the smoothed 

ice shapes had very little effect on performance results. 

The process of selecting and applying artificial material 

in order to reproduce the effects of ice roughness is 

more an art than a science. It is possible that a different 

grit size and/or distribution could produce slightly 

different results in performance degradation. 

REFERENCES

1. Addy, Jr., H. E., “Ice Accretions and Icing Effects for 

Modern Airfoils”, NASA TP 2000-210031, 

DOT/FAA/AR-99/89, April, 2000. 

2. Wright, W. B. and Rutkowski, A., “Validation 

Results for LEWICE 2.0”, NASA CR 1999-208690, 

Jan. 1999. 

3. Wright, W.B., “User Manual for the NASA Glenn Ice 

Accretion Code LEWICE Version 2.0”, NASA 

CR-1999-209409, Sept. 1999. 

4. Addy, Jr., H. E. and Chung, J. J., “A Wind Tunnel 

Study of Icing Effects on a Natural Laminar Flow 

Airfoil”, AIAA Report No. 2000-0095, Jan., 2000. 

5. Von Doenhoff, A.E., and Abbot, F.T., Jr., “The 

Langley Two-Dimensional Low-Turbulence Pressure 

Tunnel,” NACA TN 1283, May 1947. 

 6. McGhee, R.J., Beasley, W.D., and Foster, J.M., 

“Recent Modifications and Calibration of the Langley 

Low-Turbulence Pressure Tunnel,” NASA TP 2328, 

July 1984. 

 7. Paschal, K., Goodman, W., McGhee, R., Walker, B., 

and Wilcox, P.A., “Evaluation of Tunnel Sidewall 

Boundary-Layer-Control Systems for High-Lift 

Airfoil Testing,” AIAA Paper 91-3243, Sept. 1991. 

 8. Allen, H.J., and Vincenti, W.G., “Wall Interference in 

a Two-Dimensional-Flow Wind Tunnel, with 

Consideration of the Effect of Compressibility,”

NACA Report No. 782, 1944. 

 9. Nixon, D., Editor, Transonic Aerodynamics, Progress 

in Astronautics and Aeronautics, Vol. 81, pp. 489-

543, 1982. 

10. Reehorst, A.L. and Richter, G.P., “New Methods and 

Materials for Molding and Casting Ice Formations”, 

NASA TM No. 100126, September, 1987. 

11. Vickerman, M., Choo, Y., Braun, D., Baez, M., and 

Gnepp, S. “SmaggIce: Surface Modeling and Grid 

Generation for Iced Airfoils-Phase 1 Results”, AIAA 

Report No. 2000-0235, January 2000. 

12. McCullough, G.B., and Gault, D.E., “Examples of 

Three Representative Types of Airfoil-Section Stall at 

Low Speed,” NACA TN-2502, Sept. 1951. 

13. Abbott, I.H., and von Doenhoff, A.E., Theory of Wing 

Sections, Dover Publications, 1959, pp. 453-455. 

14. Drela, M., “XFOIL 6.6 Users Primer,” MIT Aero and 

Astro Engineering, Mar. 14, 1996. 

15. Ladson, C.L., “Effects of Independent Variation of 

Mach and Reynolds Numbers on the Low-Speed 

Aerodynamic Characteristics of the NACA 0012 

Airfoil Section,” NASA TM-4071, Oct. 1988. 

16. Anderson, J.D., Jr., Modern Compressible Flow with 

Historical Perspective, 2nd Ed., McGraw-Hill, New 

York, 1990, p. 279. 

17. Wootton, L.R., “The Effect of Compressibility on the 

Maximum Lift Coefficient of Aerofoils at Subsonic 

Airspeeds,” Journal of the Royal Aeronautical 

Society, Vol. 71, July 1967, pp. 476-486. 

18. Furlong, G.C., and Fitzpatrick, J.E., “Effects of Mach 

Number and Reynolds Number on the Maximum Lift 

Coefficient of a Wing of NACA 230-Series Airfoil 

Sections,” NACA TN-1299, May 1947. 

19. Furlong, G.C., and Fitzpatrick, J.E., “Effects of Mach 

Number up to 0.34 and Reynolds Number up to 8×106

on the Maximum Lift Coefficient of a Wing of 

NACA 66-Series Airfoil Sections,” NACA TN-2251, 

Dec. 1950. 

20. Morgan, H.L., Jr., Ferris, J.C., and McGhee, R.J., “A 

Study High-Lift Airfoils in the Langley LowTurbulence Pressure Tunnel,” NASA TM 89125, July 

1987. 

21. Broeren, A.P., Addy, H.E., Jr., and Bragg, M.B., 

“Effect of Intercycle Ice Accretions on Airfoil 

Performance,” AIAA Paper 2002-0240, Jan. 2002. 

22. Bragg, M.B., Khodadoust, A., and Spring, S.A., 

“Measurements in a Leading-Edge Separation Bubble 

due to a Simulated Airfoil Ice Accretion,” AIAA 

Journal, Vol. 30, No. 6, June, 1992, pp. 1462-1467. 
NASA/TMó2003-212124 8 

Fig. 1. NASA Langley Low Turbulence Pressure 

Tunnel20

Fig. 2. Business Jet Airfoil 

Fig. 3. Cast Ice Shapes 

Fig. 4. Smoothed Ice Shapes 

Fig. 5. Photograph of model in LTPT
NASA/TMó2003-212124 9 

α (deg.)

-8.0 -4.0 0.0 4.0 8.0 12.0 16.0 20.0

0.000

0.002

0.004

0.006

0.008

0.010

0.012

0.014

0.016

0.018

0.020

Cd

α (deg.)

-8.0 -4.0 0.0 4.0 8.0 12.0 16.0 20.0

-0.8

-0.4

0.0

0.4

0.8

1.2

1.6

-0.10

-0.06

-0.02

0.02

0.06

0.10

0.14

= 3.0×106

= 6.0×106

= 10.5×106

Re

Re

Re

Ma = 0.12

Cl Cm

Fig. 6 Effect of Reynolds number at constant 

Mach number on performance for the clean GLC305 airfoil. 

α (deg.)

-8.0 -4.0 0.0 4.0 8.0 12.0 16.0 20.0

-0.8

-0.4

0.0

0.4

0.8

1.2

1.6

-0.10

-0.06

-0.02

0.02

0.06

0.10

0.14

= 0.12

= 0.21

= 0.28

Ma

Ma

Ma

Re = 10.5×106

Cl Cm

α (deg.)

-8.0 -4.0 0.0 4.0 8.0 12.0 16.0 20.0

0.000

0.002

0.004

0.006

0.008

0.010

0.012

0.014

0.016

0.018

0.020

Cd

Fig. 7 Effect of Mach number at constant 

Reynolds number on performance for the clean 

GLC-305 airfoil. 
NASA/TMó2003-212124 10

(×106)

0.0 2.0 4.0 6.0 8.0 10.0 12.0

1.0

1.1

1.2

1.3

1.4

1.5

1.6

1.7

1.8

1.9

2.0

Present Data (GLC-305), = 0.12

Ladson Data (NACA 0012), = 0.15

Ladson Data (NACA 0012), = 0.20

Ladson Data (NACA 0012), = 0.25

Ma

Ma

Ma

Ma

Cl,max

Re

Fig. 8 Comparison of maximum lift Reynolds 

number dependence for the present study with the 

data of Ladson.15 

Ma

0.05 0.10 0.15 0.20 0.25 0.30 0.35

1.0

1.1

1.2

1.3

1.4

1.5

1.6

1.7

1.8

1.9

2.0

Present Data (GLC-305), = 10.5×106

Ladson Data (NACA 0012), = 6.0×106

Ladson Data (NACA 0012), = 9.0×106

Re

Re

Re

Cl,max

Fig. 9 Comparison of maximum lift Mach 

number dependence for the present study with the 

data of Ladson. 15 

Ma

0.05 0.10 0.15 0.20 0.25 0.30 0.35

-15.0

-14.0

-13.0

-12.0

-11.0

-10.0

-9.0

-8.0

-7.0

-6.0

Critcal Cp

GLC-305, = 10.5×106

Cp,min

Re

Fig. 10 Comparison of measured minimum 

pressure variation with the Critical Cp curve. 
NASA/TMó2003-212124 11

α (deg)

-8 -4 0 4 8 12 16

Cl

-0.8

-0.4

0.0

0.4

0.8

1.2

1.6

Clean

202 Casting

212 Casting

944 Casting

904 Casting

a) Lift 

α (deg)

-8 -4 0 4 8 12 16

Cd

0.00

0.02

0.04

0.06

0.08

0.10

0.12

0.14

Clean 

202 Casting

212 Casting

944 Casting 

904 Casting

b) Drag 

α (deg)

-8 -4 0 4 8 12 16

Cm

-0.16

-0.12

-0.08

-0.04

0.00

0.04

Clean 

202 Casting

212 Casting

944 Casting 

904 Casting

c) Pitching Moment 

Fig. 11. Effect of Ice Shapes on Aerodynamic 

Performance at Ma = 0.21 and Re =7.5×106

α (deg)

-8 -4 0 4 8 12 16

Cl

-0.8

-0.4

0.0

0.4

0.8

1.2

1.6

Re = 3.0x106

Re = 6.0x106

Re = 10.5x106

a) Lift 

α (deg)

-8 -4 0 4 8 12 16

Cd

0.02

0.04

0.06

0.08

0.10

0.12

0.14

Re = 3.0x106

Re = 6.0x106

Re = 10.5x106

b) Drag

α (deg)

-8 -4 0 4 8 12 16

Cm

-0.16

-0.12

-0.08

-0.04

0.00

0.04

Re = 3.0x106

Re = 6.0x106

Re = 10.5x106

c) Pitching Moment 

Fig. 12. Reynolds Number Effects on 22.5-minute 

Glaze Ice Shape (944 casting) at Ma = 0.12 
NASA/TMó2003-212124 12

α (deg)

-8 -4 0 4 8 12 16

Cl

-0.8

-0.4

0.0

0.4

0.8

1.2

1.6

Re = 3.5x106

Re = 6.0x106

Re = 10.5x106

a) Lift 

α (deg)

-8 -4 0 4 8 12 16

Cd

0.00

0.02

0.04

0.06

0.08

0.10

0.12

0.14

Re = 3.5x106

Re = 6.0x106

Re = 10.5x106

b) Drag 

α (deg)

-8 -4 0 4 8 12 16

Cm

-0.16

-0.12

-0.08

-0.04

0.00

0.04

Re = 3.5x106

Re = 6.0x106

Re = 10.5x106

c) Pitching Moment 

Fig. 13. Reynolds Number Effects on Two-minute 

Glaze Ice Shape (202 casting) at Ma = 0.12 

α (deg)

-8 -4 0 4 8 12 16

Cl

-0.8

-0.4

0.0

0.4

0.8

1.2

1.6

Re = 3.5x106

Re = 6.0x106

Re = 10.5x106

a) Lift 

α (deg)

-8 -4 0 4 8 12 16

Cd

0.00

0.02

0.04

0.06

0.08

0.10

0.12

0.14

Re = 3.5x106

Re = 6.0x106

Re = 10.5x106

b) Drag 

α (deg)

-8 -4 0 4 8 12 16

Cm

-0.16

-0.12

-0.08

-0.04

0.00

0.04

Re = 3.5x106

Re = 6.0x106

Re = 10.5x106

c) Pitching Moment 

Fig. 14. Reynolds Number Effects on 16.7-minute 

rime ice shape (212 casting) at Mach Number = 0.12 
NASA/TMó2003-212124 13

α (deg)

-8 -4 0 4 8 12 16

Cl

-0.8

-0.4

0.0

0.4

0.8

1.2

1.6

Ma = 0.12

Ma = 0.21

Ma = 0.28

a) Lift 

α (deg)

-8 -4 0 4 8 12 16

Cd

0.00

0.02

0.04

0.06

0.08

0.10

0.12

0.14

Ma = 0.12

Ma = 0.21

Ma = 0.28

b) Drag 

α (deg)

-8 -4 0 4 8 12 16

Cm

-0.16

-0.12

-0.08

-0.04

0.00

0.04

Ma = 0.12

Ma = 0.21

Ma = 0.28

c) Pitching Moment 

Fig. 15. Mach Number Effect on 22.5-minute Glaze 

Ice Shape (944 casting) at Reynolds Number = 

10.5×106

α (deg)

-8 -4 0 4 8 12 16

Cl

-0.8

-0.4

0.0

0.4

0.8

1.2

1.6

Ma = 0.12

Ma = 0.21

Ma = 0.28

a) Lift 

α (deg)

-8 -4 0 4 8 12 16

Cd

0.00

0.02

0.04

0.06

0.08

0.10

0.12

0.14

Ma = 0.12

Ma = 0.21

Ma = 0.28

b) Drag 

α (deg)

-8 -4 0 4 8 12 16

Cm

-0.16

-0.12

-0.08

-0.04

0.00

0.04

Ma = 0.12

Ma = 0.21

Ma = 0.28

c) Pitching Moment 

Fig. 16. Mach Number Effect on Two-minute Glaze 

Ice Shape (202 casting) at Reynolds Number = 

10.5×106
NASA/TMó2003-212124 14

α (deg)

-8 -4 0 4 8 12 16

Cl

-0.8

-0.4

0.0

0.4

0.8

1.2

1.6

Ma = 0.12

Ma = 0.21

Ma = 0.28

a) Lift 

α (deg)

-8 -4 0 4 8 12 16

Cd

0.00

0.02

0.04

0.06

0.08

0.10

0.12

0.14

Ma = 0.12

Ma = 0.21

Ma = 0.28

b) Drag 

α (deg)

-8 -4 0 4 8 12 16

Cm

-0.16

-0.12

-0.08

-0.04

0.00

0.04

Ma = 0.12

Ma = 0.21

Ma = 0.28

c) Pitching Moment 

Fig. 17. Mach Number Effect on 16.7-minute Rime 

Ice Shape(212 casting) at Reynolds Number = 

10.5×106

Fig. 18. Effect of Reynolds number on maximum lift 

for various GLC-305 airfoil configurations. 

Fig. 19. Effect of Mach number on maximum lift for 

various GLC-305 airfoil configurations. 

Ma

0.05 0.10 0.15 0.20 0.25 0.30 0.35

0.0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

1.6

Clean

Ice shape 202 casting

Ice shape 212 casting

Ice shape 212 smoothed

Ice shape 944 casting

Ice shape 944 smoothed

Re

Cl,max

= 10.5×106

(×106)

0.0 2.0 4.0 6.0 8.0 10.0 12.0

0.0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

1.6

Clean

Ice shape 202 casting

Ice shape 212 casting

Ice shape 212 smoothed

Ice shape 944 casting

Ice shape 944 smoothed

Ma

Cl,max

= 0.12

Re
NASA/TMó2003-212124 15

Fig. 20. Reynolds number effect on pressure 

distributions for the GLC-305 airfoil with smoothed 

ice shape 944 at α = 4 deg. 

Fig. 21. Mach number effect on pressure 

distributions for the GLC-305 airfoil with smoothed 

ice shape 944 at α = 4 deg. 

α (deg)

-8 -4 0 4 8 12 16

Cl

-0.8

-0.4

0.0

0.4

0.8

1.2

1.6

3D Casting

2D Smoothed

a) Lift 

α (deg)

-8 -4 0 4 8 12 16

Cd

0.00

0.02

0.04

0.06

0.08

0.10

0.12

0.14

3D Casting

2D Smoothed

b) Drag 

α (deg)

-8 -4 0 4 8 12 16

Cm

-0.16

-0.12

-0.08

-0.04

0.00

0.04

3D Casting

2D Smoothed

c) Pitching Moment 

Fig. 22. Casting vs. Smoothed: 22.5-minute Glaze 

Ice Shape 944 at Ma = 0.21 and Reynolds Number = 

7.5×106

x/c

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

-2.0

-1.5

-1.0

-0.5

0.0

0.5

1.0

= 3.5×106, = 0.12

= 10.5×106

, = 0.12

Cp

Re Ma

Re Ma

x/c

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

-2.0

-1.5

-1.0

-0.5

0.0

0.5

1.0

= 10.5×106, = 0.12

= 10.5×106

, = 0.28

Cp

Re Ma

Re Ma
NASA/TMó2003-212124 16

α (deg)

-8 -4 0 4 8 12 16

Cl

-0.8

-0.4

0.0

0.4

0.8

1.2

1.6

3D Casting

2D Smoothed

a) Lift 

α (deg)

-8 -4 0 4 8 12 16

Cd

0.00

0.02

0.04

0.06

0.08

0.10

0.12

0.14

3D Casting

2D Smoothed

b) Drag 

α (deg)

-8 -4 0 4 8 12 16

Cm

-0.16

-0.12

-0.08

-0.04

0.00

0.04

3D Casting

2D Smoothed

c) Pitching Moment 

Fig. 23. Casting vs. Smoothed: 16.7-minute Rime 

Ice Shape 212 at Ma = 0.21 and Reynolds Number = 

7.5×106

α (deg)

-8 -4 0 4 8 12 16

Cl

-0.8

-0.4

0.0

0.4

0.8

1.2

1.6

2D Smoothed

2D Smoothed + Grit

a) Lift 

α (deg)

-8 -4 0 4 8 12 16

Cd

0.00

0.02

0.04

0.06

0.08

0.10

0.12

0.14

2D Smoothed

2D Smoothed + Grit

b) Drag 

α (deg)

-8 -4 0 4 8 12 16

Cm

-0.16

-0.12

-0.08

-0.04

0.00

0.04

2D Smoothed

2D Smoothed + Grit

c) Pitching Moment 

Fig. 24. Smoothed vs. Grit: 22.5-minute Glaze Ice 

Shape 944 at Ma = 0.21 and Reynolds Number = 

7.5×106
NASA/TMó2003-212124 17

α (deg)

-8 -4 0 4 8 12 16

Cl

-0.8

-0.4

0.0

0.4

0.8

1.2

1.6

2D Smoothed

2D Smoothed + Grit

a) Lift 

α (deg)

-8 -4 0 4 8 12 16

Cd

0.00

0.02

0.04

0.06

0.08

0.10

0.12

0.14

2D Smoothed

2D Smoothed + Grit

b) Drag 

α (deg)

-8 -4 0 4 8 12 16

Cm

-0.16

-0.12

-0.08

-0.04

0.00

0.04

2D Smoothed

2D Smoothed + Grit

c) Pitching Moment 

Fig. 25. Smoothed vs. Grit: 16.7-minute Rime Ice 

Shape 212 at Ma = 0.21 and Reynolds Number = 

7.5×106
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Unclassified Unclassified

Technical Memorandum

Unclassified

National Aeronautics and Space Administration

John H. Glenn Research Center at Lewis Field

Cleveland, Ohio 44135–3191

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National Aeronautics and Space Administration

Washington, DC 20546–0001

Available electronically at http://gltrs.grc.nasa.gov

February 2003

NASA TM—2003-212124

AIAA–2003–0727

E–13776

WBS–22–708–20–02

23

A Wind Tunnel Study of Icing Effects on a Business Jet Airfoil

Harold E. Addy, Jr., Andy P. Broeren, Joseph G. Zoeckler, and Sam Lee

Aircraft icing; Airfoil; Aerodynamic performance

Unclassified -Unlimited

Subject Category: 02 Distribution: Nonstandard

Prepared for the 41st Aerospace Sciences Meeting and Exhibit sponsored by the American Institute of Aeronautics and

Astronautics, Reno, Nevada, January 6–9, 2003. Harold E. Addy, Jr. and Joseph G. Zoeckler, NASA Glenn Research

Center; Andy P. Broeren, University of Illinois at Urbana-Champaign, Urbana, Illinois 61801; and Sam Lee, QSS Group,

Inc., Cleveland, Ohio 44135. Responsible person, Harold E. Addy, Jr., organization code 5840, 216–977–7467.

Aerodynamic wind tunnel tests were conducted to study the effects of various ice accretions on the aerodynamic

performance of a 36-inch chord, two-dimensional business jet airfoil. Eight different ice shape configurations were

tested. Four were castings made from molds of ice shapes accreted in an icing wind tunnel. Two were made using

computationally smoothed tracings of two of the ice shapes accreted in the icing tunnel. These smoothed profiles were

then extended in the spanwise direction to form a two-dimensional ice shape. The final two configurations were formed

by applying grit to the smoothed ice shapes. The ice shapes resulted in as much as 48 percent reduction in maximum lift

coefficient from that of the clean airfoil. Large increases in drag and changes in pitching moment were also observed.

The castings and their corresponding smoothed counterparts yielded similar results. Little change in performance was

observed with the addition of grit to the smoothed ice shapes. Changes in the Reynolds number (from 3×106 to

10.5×106) and Mach number (from 0.12 to 0.28) did not significantly affect the iced-airfoil performance coefficients.