Harold E. Addy, Jr.

Glenn Research Center, Cleveland, Ohio

Sam Lee

ASRC Aerospace Corporation, Cleveland, Ohio

Icing Encounter Duration Sensitivity Study

NASA/TM—2011-216367

March 2011

AIAA–2009–4263
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Harold E. Addy, Jr.

Glenn Research Center, Cleveland, Ohio

Sam Lee

ASRC Aerospace Corporation, Cleveland, Ohio

Icing Encounter Duration Sensitivity Study

NASA/TM—2011-216367

March 2011

AIAA–2009–4263

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

First Atmospheric and Space Environments Conference

sponsored by the American Institute of Aeronautics and Astronautics

San Antonio, Texas, June 22–25, 2009
Acknowledgments

The authors would like to thank the crew of the IRT for providing expert assistance throughout the tests, Julius Mirecki

for setting up the measurement systems and providing support for the data acquisition, and to Andy Broeren for

providing key insight and suggestions during data analysis.

Available from

NASA Center for Aerospace Information

7115 Standard Drive

Hanover, MD 21076–1320

National Technical Information Service

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

Level of Review: This material has been technically reviewed by technical management. 
NASA/TM—2011-216367 1 

Icing Encounter Duration Sensitivity Study 

Harold E. Addy, Jr. 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Sam Lee 

ASRC Aerospace Corporation 

Cleveland, Ohio 44135 

Abstract 

This paper describes a study performed to investigate how aerodynamic performance degradation 

progresses with time throughout an exposure to icing conditions. It is one of the first documented studies 

of the effects of ice contamination on aerodynamic performance at various points in time throughout an 

icing encounter. Both a 1.5 and a 6 ft chord, two-dimensional, NACA-23012 airfoils were subjected to 

icing conditions in the NASA Icing Research Tunnel (IRT) for varying lengths of time. At the end of each 

run, lift, drag, and pitching moment measurements were made. Measurements with the 1.5 ft chord model 

showed that maximum lift and pitching moment degraded more rapidly early in the exposure and 

degraded more slowly as time progressed. Drag for the 1.5 ft chord model degraded more linearly with 

time, although drag for very short exposure durations was slightly higher than expected. Only drag 

measurements were made with the 6 ft chord airfoil. Here, drag for the long exposures was higher than 

expected. Novel comparison of drag measurements versus an icing scaling parameter, accumulation 

parameter times collection efficiency was used to compare the data from the two different size models. 

The comparisons provided a means of assessing the level of fidelity needed for accurate icing simulation. 

Nomenclature 

Ac accumulation parameter 

AoA angle of attack, deg 

beta collection efficiency 

c chord 

cd drag coefficient 

cd0 drag coefficient, clean airfoil 

Cl lift coefficient 

Clm maximum lift coefficient 

Clm0 maximum lift coefficient, clean airfoil 

Cm pitching moment coefficient 

F Fahrenheit k = ice height 

LWC liquid water content, grams/cubic meter 

MVD median volumetric diameter, micrometers 

Re Reynolds number 

V airspeed, knots 
NASA/TM—2011-216367 2 

Introduction 

Airfoil and aircraft performance degradation due to inflight icing is a widely known aviation hazard. 

Over the last two decades, studies (Refs. 1 to 17) have been conducted to investigate and measure the 

effect ice accretions have on the aerodynamic performance of airfoils and wings. Various aspects of the 

problem have been considered including the type of ice, the fidelity of the ice shapes, and the type of 

airfoil. Little has been done, however, to study how aerodynamic performance degrades with time as the 

ice initially begins to accrete on an airfoil and then grows as the icing encounter continues. 

One study that did investigate aerodynamic performance degradation with time as ice continued to 

accrete was conducted by Potapczuk and Berkowitz in 1990 (Ref. 17). Using a two-dimensional multielement airfoil model mounted on a force balance in the NASA Glenn Research Center’s Icing Research 

Tunnel (IRT), measurements of lift, drag, and pitching moment were made as ice accreted on the model. 

Large degradations in performance were measured. The degradations, however, were typically not linear 

with time. In general, the initial ice accretions contributed more to the overall degradation than ice 

accreted later during the icing encounter. 

The Potapczuk and Berkowitz study, however, did not attempt to measure icing effect on maximum 

lift coefficient or stall angle throughout the accretion. These measurements were made only at the end of 

the allotted ice accretion time. Another aspect of the Potapczuk and Berkowitz study was that it was 

performed on a subscale model. Because ice accretes differently on a subscale model than on a full scale 

model, there can be different aerodynamic effects. The considerable resources required for a full scale 

investigation were unavailable for that study, although the results were still very valuable. 

The investigation described in this paper was designed to address the outstanding questions left by the 

Potapczuk and Berkowitz study. Airfoil aerodynamic performance measurements were made after various 

lengths of exposure to icing conditions. Both subscale and full scale, two-dimensional models were 

tested. The icing conditions remained the same throughout the test for each model, only the length of 

exposure was varied. For the iced subscale model, lift and pitching moment measurements were made at 

multiple angles of attack from zero through maximum lift. Drag measurements were made at selected 

angles of attack for both models with ice accreted. 

Test Description 

This study included two test campaigns in the NASA Glenn Icing Research Tunnel. The first was 

conducted using a two-dimensional, subscale, 1.5 ft chord, NACA-23012 model. The second campaign 

used a full-scale, 6 ft chord version of the same airfoil. The models had been used in prior ice accretion 

and iced-aerodynamic performance studies, so existing databases could be leveraged to verify and 

corroborate data collected in this study. The models were exposed to selected icing conditions in the IRT 

for several different periods of time. At the conclusion of each exposure period, aerodynamic 

performance measurements were made and then the ice accretions were documented using ice tracings, 

photographs, and ice depth measurements. 

First Campaign 

The first test used the 1.5 ft chord NACA-23012 model. There were two reasons for testing this 

model in the first campaign. First, its size, in comparison with the dimensions of the test section of the 

IRT, allowed lift and pitching moment data to be taken free of large wall and blockage interference 

effects. Secondly, since high-quality aerodynamic performance measurements had been made in a lowspeed aerodynamic tunnel using castings of ice shapes accreted at some of the same conditions as those in 

this study, the quality of the performance measurements in the IRT could be verified. 

A photograph of the 1.5 ft chord model mounted vertically in the IRT is shown in Figure 1. It was 

made from solid aluminum and had a row of pressure taps located about 18 in. above the floor of the 

tunnel. These pressure taps were used to align the clean model to the airflow and obtain clean model lift 
NASA/TM—2011-216367 3 

and pitching moment coefficients. Because the taps around the leading edge of the models iced over 

during exposure to the icing cloud, the pressure measurements were not used to make lift and pitching 

moment calculations after ice was accreted. 

The model was mounted on an external force balance located just below the floor and above the 

ceiling of the tunnel. The measurements of lift and pitching moment with ice accreted on the model were 

made using this force balance system. Conventional two-dimensional wind tunnel model corrections 

(Ref. 18) were applied to all the performance data. Drag measurements were made using a traversing 

wake survey system mounted on the floor downstream of the model. 

Icing Research Tunnel 

The IRT is an atmospheric wind tunnel with a test section 6 ft high by 9 ft wide and 20 ft long. It is 

capable of airspeeds up to 400 mph (test section empty), Mach numbers to about 0.45, and Reynolds 

numbers of three million per foot. Air temperatures can be cooled to –20 F. 

It can produce an icing cloud with droplet sizes and liquid water contents similar to those prescribed 

in the FAA’s CFR Appendix C aircraft certification icing cloud specifications. It can also produce 

supercooled large droplets (SLD) over a range of droplet sizes and liquid water contents. 

The ability of the IRT to repeat ice shapes for a given set of icing conditions is well established 

(Ref. 22). Tracings were made of select repeat conditions to ensure measurements were being made with 

like ice accretions. 
NASA/TM—2011-216367 4 

A difficulty with making aerodynamic performance measurements in an icing wind tunnel is that these 

tunnels typically have high freestream turbulence levels in comparison with dry, aerodynamic wind tunnels. 

The cloud generating equipment present in the tunnel is one of the reasons for this. Moreover, it is not 

unusual for equipment to be added to an icing tunnel specifically to increase air turbulence to make the icing 

cloud more uniform. A uniform icing cloud is often more important to icing tests than elevated turbulence. 

Finally, to prevent the water droplet generating equipment (i.e., water spray systems) from freezing, air is 

continuously blown through the droplet-generating nozzles, thereby generating additional air turbulence. 

Turbulence in the IRT has been measured at about 1.0 percent with spray bar air activated (Ref. 19). 

Test Setup 

The 1.5 ft chord model was mounted vertically on the turntable and force balance in the test section of 

the IRT and spanned the 6 ft from floor to ceiling. A traversing drag wake survey system was mounted to 

the floor of the tunnel about two chord lengths downstream of the model’s trailing edge. 

The icing condition selected was a glaze ice condition from the CFR Appendix C envelope. Glaze ice 

accreted on the leading edge of an airfoil typically produces a high degree of performance degradation, so 

a study of how this degradation changes with time of exposure is of particular interest. For the 1.5 ft 

chord model, a cloud with droplets of 15 μm in median volumetric diameter (MVD) and with a liquid 

water content (LWC) of 0.75 gm3

 was chosen. The models were set at an angle-of-attack (AOA) of 2. 

The airspeed was 200 knots and the air total air temperature was 28 F. This condition had been used in a 

previous study where aerodynamic performance measurements were made in a dry wind tunnel using 

castings. This allowed iced performance measurements made in the IRT to be compared with similar 

measurements in another tunnel. The cloud covered most, but not all, of the model span. The ice 

accretions started at about 6 in. from the ceiling and ended about 3 in. from the floor of the tunnel. 

Another difficulty discovered during the tests was that at the 200 knot airspeed at which the ice was 

accreted, the wake survey probe tip tended to ice up due to a very faint cloud of very small (<10 μm 

diameter) water droplets circulating with the air in the tunnel. These droplets are so small and constitute 

such a low LWC that they do not affect most tests in the IRT. However, they do form small accretions on 

smaller objects with sharp edges where the collection efficiency is very high. The tip of the wake survey 

probe is heated to prevent ice buildup, but, at 200 knots and 28 F, the convective cooling of the tip is 

sufficient to permit ice accretion. Lowering the airspeed to 150 knots allowed the wake survey tip to 

remain warm enough to prevent any ice accretion. Therefore, ice was first accreted on the model at the 

stated conditions, and then the tunnel airspeed was lowered to 150 knots where the aerodynamic 

performance data were acquired. This was done for both the first and second campaigns of this study. The 

Mach number at this speed and temperature was 0.235. The Reynolds numbers were 2.6 and 10.4 million 

for the 1.5 and 6 ft chord models, respectively. 

A final challenge for these tests was that once the cloud in the IRT was shut off, the accreted ice 

would slowly begin to sublimate or erode, or both. This meant that if it took a long time to acquire the 

performance data, the ice accretion at the end of the data acquisition was different than at the beginning. 

Since taking a complete set of performance data could be time consuming, ice accretions were often 

repeated acquiring partial data sets each time, in order to acquire a complete set of data. 

Force balance data were often collected at four to six angles-of-attack at a time for each ice accretion. 

Data were taken mostly around stall. Wake survey data were taken at four angles-of-attack, 0, 2, 4, and 

6. It usually required two ice accretions to acquire the wake survey data for a particular icing exposure 

time. 

Measurement Errors 

An error analysis of the algorithm and the pressure instrumentation used to calculate the drag 

coefficients show that the uncertainty for these values is 12 drag counts where one drag count is an 

increment of 0.0001 in drag coefficient. While relatively high for clean airfoil drag measurements, it is a 

reasonable level for ice-contaminated airfoils where the drag coefficients are typically much higher. 
NASA/TM—2011-216367 5 

A static calibration of the force balance has shown the accuracy of this measurement system to be 

3 percent in lift and 10 percent in pitching moment. It has been noted (Ref. 20), however, that for twodimensional wind tunnel tests where a constant chord model completely spans the test section, the 

spanwise distribution of lift caused by the tunnel walls diminishes lift by, typically, 10 to 15 percent at 

maximum lift coefficient. This has been observed in the IRT (Ref. 7). 

A further potential complication exists when using the force balance in the icing tunnel. As was 

mentioned previously, the icing cloud does not cover the entire span of the model, but the force balance 

measures the forces over the entire model. Therefore, some error is introduced into the assumption of 

infinite span, as is the case in full span, two-dimensional model testing. 

An error analysis (Ref. 7) of the data measurements indicates that the lift coefficients calculated from 

the surface pressures have an uncertainty of approximately 2 percent up to the maximum lift coefficient 

and 4 percent for pitching moment. 

First Campaign Results 

Because aerodynamic performance measurements are not often made in an icing wing tunnel, in 

addition to establishing system component error, the repeatability of the measurements was documented. 

Figure 2 shows lift and pitching moment coefficients for the clean model as measured by the force 

balance on four different days of testing. As can be seen, the repeatability was very good. Similar 

repeatability curves for the pressure tap measurements are shown in Figure 3. Data from Run 210 were 

not shown because water had contaminated the pressure lines from the taps to the transducers. The water 

was subsequently cleared. 

Repeat runs were not made with the wake survey for the clean model due primarily to the large 

investment of test time required to run a survey. However, repeat wake surveys were run for the iced 

airfoil. These data are shown in Figure 4. Again, the repeatability is good. 

Also to increase confidence in the data, measurements were compared with data measured at the 

University of Illinois at Urbana-Champaign’s Low Speed Wind Tunnel (LSWT). Figure 5 shows a 

comparison of lift and pitching moment coefficients as measured by both the force balance and the 

surface pressures in the IRT with those measured in the LSWT using the surface pressure method. 

 
NASA/TM—2011-216367 6 

 

The models in both tests were 1.5 ft chord, NACA-23012 models. However, the model used in the 

LSWT had a span of 34 in., rather than the 6 ft span of the IRT model. Also, the data were collected at 

differing Reynolds numbers. The LSWT data were collected at a Re = 1.8 million while the IRT data 

were collected at Re = 2.35 million. The difference in Reynolds number explains the lower maximum lift 

coefficient and earlier stall angle for the LSWT test. This is illustrated in Figure 6 where data from tests 

conducted by the UIUC of a NACA-23012 airfoil in NASA Langley’s Low Turbulence Pressure Tunnel 

(LTPT) is compared with data from the current study in the IRT. The pressure capability of the LTPT 

allowed independent variation of Mach and Reynolds numbers. The IRT data were acquired at a Reynolds 

number of 2.6 million, while the data shown from the LTPT are from Reynolds numbers of 2.0 and 

3.5 million. As expected, the IRT data falls in between the data from the LTPT. 

The difference in the lift coefficients near maximum lift between the force balance and surface 

pressure measurements in the IRT is likely due to airflow at the floor and ceiling of the IRT which causes 

a decrease in the lifting force on the airfoil. This decrease is measured by the force balance but not by the 

surface pressure taps which measure only section lift at a station 18 in. above the floor. The difference in 

pitching moment coefficients between the force balance and surface pressure measurements in the IRT is 

likely due to the inaccuracy of the force balance measurement. To obtain the pitching moment, the force 

balance must discern a small difference in two much larger forces and then multiply the difference by the 

moment arm. The error band becomes relatively large in these circumstances. Nonetheless, the trends are 

correct. 

A comparison of clean model drag wake survey results for the IRT and LSWT tests is shown in 

Figure 7. Drag coefficients measured in the IRT are slightly higher likely due to the higher turbulence 

levels in the IRT. The higher Reynolds number at which the IRT data were taken may also contribute to 

the higher IRT results. 

Iced Airfoil Results 

For the 1.5 ft chord model, aerodynamic performance measurements were made after icing exposure 

times of 0.25, 0.5, 1.0, 2.5, and 5.0 min. The ice accreted during the latter four of these exposure times is 

shown in the ice tracings of Figures 8(a) to (d). No tracing or ice depth measurement was made of the 

0.25 min exposure ice accretion because the amount of ice was too small. 
NASA/TM—2011-216367 7 

 
NASA/TM—2011-216367 8 

 

Force balance measurements made for these ice accretions are shown in Figure 9. Even the very small 

amount of ice accreted during the 0.25 min exposure had a dramatic effect on maximum lift coefficient 

and stall angle. Note, however, that the leading edge stall exhibited by the airfoil for this small amount of 

ice is the same as for the clean airfoil. At 0.5 min of ice accretion, maximum lift coefficient and stall 

angle are further diminished, but a different type of stall is indicated by the less-abrupt decrease in the lift 

curve. At the Reynolds number of 2.6 million at which these data was obtained, it is likely that a thin 

airfoil-type stall occurred. The effects of glaze ice on lift and pitching moment for this airfoil are 

dramatic. The 5 min. ice accretion reduced maximum lift to 38 percent of the clean airfoil’s value. 

The pitching moment curves for each of these ice accretions are also shown in Figure 9. These data 

also reflect the dramatic effect such ice accretions have on the aerodynamic performance of this model. 

Note here also, that the very short icing exposure time of only 0.25 min causes the pitching moment to go 

strongly negative at a much lower angle of attack than for the clean airfoil. 

The drag data for these ice accretions also show dramatic aerodynamic performance degradation as 

shown in Figure 10. Again, a relatively large increase in drag was caused by the small, 0.25 min ice 

accretion. 

As was mentioned earlier, some of the ice accretions matched ice accretions from a previous test 

where ice castings were made and subsequently tested in the dry, aerodynamic, LSWT at UIUC. The 

performance data from those tests are compared with the IRT data in Figures 11 and 12. The lift 

coefficients for the 0.5 min ice accretion and casting agree very well even though the Reynolds numbers 

for the two cases are different. In several two-dimensional airfoil tests with leading edge ice, however, it 

has been shown that Reynolds number effects are negated by the presence of the ice on the leading edge 

of the model (Refs. 5, 6, and 12). There is some difference in the lift coefficients for the 5 min ice shape 

around maximum lift with the IRT values showing greater degradation. However, the lift curves are very 

similar in shape indicating a similar stall behavior. 

The drag coefficients agree fairly well for the 0.5 min ice accretion at IRT and the corresponding 

casting at LSWT. Similar to the lift coefficients results for the 5.0 min ice shapes, the IRT accretion 

shows greater performance degradation in comparison with the LSWT results. 
NASA/TM—2011-216367 9 

 

 

Another way to look at performance degradation with exposure time is to plot maximum lift 

coefficient and drag coefficient, each normalized by its appropriate uncontaminated model value, as a 

function of exposure time. Figure 13 shows this plot for maximum lift coefficient normalized by the 

maximum lift coefficient for the clean model. The curve is roughly asymptotic, indicating diminishing 

maximum lift coefficient degradation with exposure time. 

Figure 14 shows drag coefficients normalized by the appropriate clean model drag coefficient as a 

function of exposure time for each of the four angles of attack at which wake survey measurements were 

taken. For each angle of attack, the drag coefficients form approximately straight lines. In contrast to lift 
NASA/TM—2011-216367 10 

degradation, the drag degrades linearly with exposure time. There is, however, a modest, but noticeable 

bump up in the drag versus exposure time curves for the very short exposure time of 0.25 min. This short 

exposure time is not long enough to allow the airfoil’s surface to become thoroughly wetted. The 

appearance of the ice accreted is patchy and somewhat rough and does not completely cover the leading 

edge of the airfoil. The patchy roughness of the ice is likely more a function of the water surface tension 

and airfoil surface character and is less dependent on the size of the model. This raises the question of 

whether the initial ice roughness on the leading edge of the model is disproportionately large for a small 

subscale model. 

Also shown in Figure 14 are drag data from one of the Potapczuk and Berkowitz tests. Although the 

model, icing conditions, and ice accretion are different, the trend is similar. 

Second Campaign Results 

A second test campaign was conducted in the IRT using a 6 ft chord, two-dimensional NACA-23012 

airfoil in order to obtain reference data for the small scale, 1.5 ft chord model tests. The performance 

degradation due to the short icing exposures of the 1.5 ft chord model seemed inordinately high. Running 

tests with a 6 ft chord model could help determine whether the initial ice accretion was creating an ice 

roughness that was disproportionately large for the size of the 1.5 ft chord model. 

Like the 1.5 ft chord model, the 6 ft chord model was also machined from solid aluminum and had a 

chord-wise row of pressure taps located 18 in. above the tunnel floor on the vertically mounted model. A 

photo of the model installed in the test section of the IRT is shown in Figure 15. 
NASA/TM—2011-216367 11 

Aerodynamic performance measurements were limited with the 6 ft chord model for several reasons. 

First, the model could not be mounted on the IRT’s external force balance due to the manner in which its 

attachment points were constructed. Secondly, due to the large blockage created by the model in the 

tunnel, it could not be rotated through stall. Finally, as was the case for the 1.5 ft chord model, when ice 

was accreted, the leading edge surface pressure taps were covered, preventing lift and pitching moment 

calculations. Therefore, only drag measurements were made at select low angles of attack for the clean 

and iced model and lift and pitching moment for the clean model. Drag was measured using the same 

wake survey system as used in the first campaign. 

The icing conditions selected for the second campaign were different from those used in the first 

campaign due primarily to the difference in collection efficiencies of the two models. The icing 

conditions for the two campaigns are shown in Table 1 along with the collection efficiencies for the two 

models and the accumulation parameters for the different cases. 

TABLE 1.—ICING CONDITIONS FOR THE IRT CAMPAIGNS 

Two-dimensional NACA‐23012 

Chord MVD LWC V AoA Exposure 

time 

Accumulation 

parameter 

Collection 

efficiency 

Ac‐beta Similar cases 

ft m–6

 g/m3

 kt deg min Ac beta 

1.5 15.4 0.75 200 2 0.25 0.087 0.701 0.061 A 

1.5 15.4 0.75 200 2 0.5 0.17 0.701 0.12 B 

1.5 15.4 0.75 200 2 1 0.35 0.701 0.25 C 

1.5 15.4 0.75 200 2 2.5 0.87 0.701 0.61 D 

1.5 15.4 0.75 200 2 5 1.7 0.701 1.2 

 

6 20 0.5 200 2 1 0.058 0.511 0.030 

6 20 0.5 200 2 2 0.12 0.511 0.060 A 

6 20 0.5 200 2 5 0.29 0.511 0.15 B 

6 20 0.5 200 2 7.5 0.44 0.511 0.22 C 

6 20 0.5 200 2 10 0.58 0.511 0.30 

6 20 0.5 200 2 15 0.87 0.511 0.45 

6 20 0.5 200 2 22.5 1.3 0.511 0.67 D 

Collection efficiencies and accumulation parameters (Ref. 21), along with their product, Ac-beta, 

were calculated for the different cases to obtain geometrically similar ice shapes for the two different size 

models. The cases A, B, C, and D in the table are pairs of geometrically similar ice accretions. Tracings of 

the ice shapes for the 6 ft chord model test are shown in Figures 16(a) to (d). The ice accretions for the 1 

and 2 min cases for the 6 ft chord model were too small to trace. 

To raise confidence in the aerodynamic measurements in the IRT, the IRT data were compared with 

data taken in a dry aerodynamic tunnel in another program. In an iced aerodynamic simulation program 

conducted jointly with NASA, ONERA, and UIUC participation, a two-dimensional, 6 ft chord, NACA23012 airfoil was tested both clean and with cast ice shapes in ONERA’s F1 tunnel at Fauga-Mauzac. 

Drag data from clean models are compared in Figure 17, while drag data from the ice-contaminated 

models are compared in Figure 18. 

For the clean model, the drag coefficients measured in the IRT are high compared with the dry 

aerodynamic tunnel data. There were some differences in the Reynolds and Mach numbers at which the 

data was taken in the two tunnels (Re = 9.1 million and M = 0.20 in F1 versus Re = 10.4 million and 

M = 0.235 in IRT), but the primary difference was the higher turbulence levels in the IRT. 

The ice accretion case that matched an ice casting tested at the F1was for a 2 min exposure time. At 

low angles of attack, 0 and 2, the drag coefficients measured in the IRT are high. However, for angles 

of attack of 4 and 6, the drag coefficients agree well. These results might be explained by: 1) at the 

lower AoA’s for this amount of ice, the drag is predominantly due to skin friction where the higher 

turbulence level in the IRT has an influence, and 2) at the higher angles, a larger area of the airfoil is 

covered by the trailing edge separation zone and the effects due to IRT turbulence are less important. 
NASA/TM—2011-216367 12 

 
NASA/TM—2011-216367 13 

Since most of the data with the 6 ft chord model in the IRT were acquired with ice accretion times greater 

than 2 min, the drag results should be reasonable. 

The set of drag coefficient curves versus angle of attack for the various icing exposure times is shown 

in Figure 19. 

Normalizing the ice tracings and the models by chord length allows the ice shapes to be compared for 

geometric similarity. The cross plotted ice shapes for three of the four comparable Ac-beta cases are 

shown in Figures 20(a) to (c). The ice tracings for Cases B and C are very similar. The tracings for the 6 ft 

chord model do seem to be a bit rougher, but that may be an artifact of the tracing process. Tracing ice 

does result in some smoothing of the finer ice roughness. Such roughness may have been smoothed over 

when tracing the ice on the 1.5 ft chord model, but such features on the 6 ft model were large enough to 

be traced. 

However, there are some differences in the Case D ice shapes, even though, overall, the amount of ice 

is about the same and there are lower and upper surface horns on both. The upper surface horn on the 6 ft 

chord model is taller and narrower, and is angled more forward than the ice accretion for the 1.5 ft chord 

model. The lower surface horn has about the same angle, but is taller and narrower. 

Figure 21 shows the drag coefficient versus angle of attack curves plotted along with the curves from 

the first campaign for the clean airfoils as well as for the iced airfoils with similar Ac-beta values. The 

clean airfoil drag curves agree well. The Ac-beta Case B and C curves also agree well with the drag 

values from the 1.5 ft chord model. However, neither the Case A nor the Case D drag curves agree well 

with the curves from the 1.5 ft chord model. For Case A, the drag values for the 1.5 ft chord model were 

significantly higher than for the 6 ft chord model. These data confirm the suspicion mentioned earlier that 

drag for an initial ice shape on the 1.5 ft airfoil was too high and not representative of an initial ice 

accretion on a full scale airfoil. For Case D, however, the drag coefficients from the 6 ft chord model are 

much higher than those from the 1.5 ft chord model tests. This result was not expected. 

The higher than expected drag coefficients for the longer icing exposures are also evident when 

plotting the 6 ft chord model data versus exposure time as shown in Figure 22. Unlike the curves for the 

1.5 ft model as shown in Figure 14, the curves form straight lines for the shorter icing duration periods. 

However, there is a noticeable break in the slope of the curves at the icing exposure time of 10 min for the 

angles of attack of 0 and 2 and 5 min for the angles of attack of 4 and 6. 
NASA/TM—2011-216367 14 

 
NASA/TM—2011-216367 15 

 

This observation is also demonstrated in the plot of normalized drag coefficient versus Ac-beta for 

both models at 2 AoA, as shown in Figure 23. In this plot, the higher drag values for the 1.5 ft chord 

model at lower Ac-beta values are easily seen as are the much greater drag values for the 6 ft chord model 

at the higher Ac-beta values. It indicates that something about the larger ice shapes on the 6 ft chord 

model is having a significantly different effect on aerodynamic drag. 

The difference in drag for larger ice shapes between the two models is also seen in Figure 24 where 

drag values for the two airfoils are plotted versus Ac-beta, only this time at 6 angle of attack. Here, the 

difference in drag appears at Ac-beta values as low as 0.2. These data reflect the increasing importance of 

the upper surface horn to the airflow at higher angles of attack. The reader will note that, at low Ac-beta 

values, the curves for the two models agree. This suggests that the increase in drag for the smaller model 

at short icing exposure times is due to ice on the lower surface of the airfoil. 

For each ice tracing made in the IRT, an ice depth gage was used to measure the maximum thickness 

of ice on both the upper and lower airfoil surfaces. The thickness of the ice at the stagnation line was also 

measured. These measurements are shown in Table 2 along with the corresponding values normalized by 

chord length, k/c. The k/c values are also plotted as a function of Ac-beta in Figures 25 and 26. Figure 25 

shows the upper surface and stagnation line values for both models. The stagnation line values for the two 

models agree well. This is not surprising because icing scaling methods are based upon stagnation line 

similarity assumptions. The normalized upper surface maximum ice thickness values for the 6 ft chord 

model are, however, higher than those for the 1.5 ft chord model. This result is likely due to the 

complexity of the glaze ice accretion process where airspeed, air temperature, amount of water, and 

boundary layer dynamics all play a role in where the flowing water freezes as it flows downstream. Such 

details are difficult, if not impossible, to capture in icing scaling methods. 

Maximum ice thickness data for the lower surface are shown in Figure 26. The data for the two 

models match remarkably well for accretions up to an Ac-beta value of 0.30. This corresponds to the 

10 min ice accretion for the 6 ft chord model. It also corresponds to the point where the slope changes in 

the normalized drag versus icing exposure time curves for 0 and 2 AoA in Figure 22 as well as in the 

6 ft chord model normal drag versus Ac-beta curve of Figure 23. 
NASA/TM—2011-216367 16 

TABLE 2.—ICE THICKNESS DATA 

Two-dimensional NACA‐23012 

Chord Exp. 

time 

Ac‐beta Ice Thickness, in. k/c Similar cases 

ft min Upper Stagnation Lower (upper) (stag) (lower) 

1.5 0.25 0.061 ------- ------- ------- ------- ------- ------- A 

1.5 0.5 0.12 0.04 0.03 0.03 0.002 0.002 0.002 B 

1.5 1 0.25 0.11 0.05 0.12 0.0061 0.0028 0.0067 C 

1.5 2.5 0.61 0.39 0.09 0.27 0.022 0.005 0.015 D 

1.5 5 1.2 0.81 0.15 0.53 0.045 0.0083 0.029 

 

6 1 0.030 ------- ------- ------- ------- ------- ------- 

6 2 0.060 ------- ------- ------- ------- ------- ------- A 

6 5 0.15 0.42 0.12 0.25 0.0058 0.0017 0.0035 B 

6 7.5 0.22 0.64 0.17 0.44 0.0089 0.0024 0.0061 C 

6 10 0.30 0.92 0.18 0.6 0.013 0.0025 0.0083 

6 15 0.45 1.28 0.28 0.96 0.0178 0.0039 0.0133 

6 22.5 0.67 1.88 0.38 1.71 0.0261 0.0053 0.0238 D 

 
NASA/TM—2011-216367 17 

These observations indicate that the character of the ice for the longer exposure times on the 6 ft 

chord model is responsible for the higher drag values when compared with those of the 1.5 ft chord 

model. The upper surface horn on the 6 ft chord model is slightly higher, less-rounded, and has a slightly 

different position on and angle with respect to the local surface than the upper surface horn on the 1.5 ft 

chord model. Furthermore, it is interesting and important to note that in both the case of the 1.5 ft chord 

model with the initial ice accretion and in the case of the 6 ft chord airfoil with the long-exposure ice 

accretion, it is the character of the ice that results in increased drag. 

A further observation from the data of this study is that not only are features of the ice important, but 

when during the ice accretion they begin to take shape is important. The drag for the 6 ft airfoil increased 

at an accelerated rate once both upper and lower surface horns were established. 

Finally, the novel analysis of aerodynamic performance in terms of parameters typically used of icing 

scaling provides a means to assess the level of fidelity required for effective icing simulation. 

Concluding Remarks 

The primary objective of this study was to investigate how aerodynamic performance degrades as ice 

accretes on an airfoil. Existing two-dimensional airfoil models were used which had an established 

background of ice-contaminated aerodynamic performance data. Glaze icing conditions were selected 

which had shown to be particularly detrimental to aerodynamic performance. 

The two models used in the study were both NACA-23012 airfoils. The difference between them was 

size. The smaller model had a chord of 1.5 ft while the larger one had a chord of 6 ft; typical of airfoils of 

this family in use on aircraft. The smaller model was used because it could be rotated through stall in the 

NASA Icing Research Tunnel. The larger model could not without inducing undue blockage in the tunnel. 

Both models had a floor-to-ceiling span of 6 ft. 

Results from the small model testing showed that maximum lift coefficient and stall angle degraded 

in a diminishing manner as time exposed to icing conditions increased. In general, drag coefficient 

increased linearly with time exposed to icing conditions, with the exception of short exposure times where 

increases in drag coefficient were more pronounced. 

Only aerodynamic drag results were available from the tests with the large, 6 ft chord model. For 

these tests, the large model could not be mounted to the IRT’s external force balance and ice covered 

surface pressure taps needed to make accurate performance calculations. Due to the large model’s lower 

collection efficiency, slight changes were made to cloud droplet size, liquid water content, and exposure 

times to result in geometrically similar ice accretions as in the smaller model test. 

Results from the large model test showed that drag initially increases linearly with icing exposure 

time but, for longer ice accretion times, it increases more rapidly. The cause of this behavior appeared to 

be the establishment of relatively taller and narrower ice horns on the upper and lower surfaces of the 

airfoil at certain points during exposure to icing conditions. These ice features were different from those 

which formed on the small scale airfoil and were likely the reason for differences in drag coefficients 

measured on the two airfoils even though the ice shapes were relatively similar. Furthermore, the study 

emphasized that not only are particular ice features important, but also the point in time during an icing 

encounter when they develop is important. 

In order to compare the data between the subscale and full scale airfoils, aerodynamic performance 

data were assessed using icing scaling parameters. The correlations provide a novel manner in which 

icing data can be analyzed. Moreover, it provides a means by which the fidelity of icing simulations can 

be assessed; a goal that has remained elusive in icing technology. 
NASA/TM—2011-216367 18 

References 

1. Bragg, M., Broeren, A., Addy, H., Potapczuk, M., Guffond, D., Montreul, E., “Airfoil Ice-Accretion 

Aerodynamics Simulation,” AIAA–2007–0085, 2007. 

2. Papadakis, M., Yeong, H.W., Vargas, M., Potapczuk, M.P., “Aerodynamic Performance of a Swept 

Wing with Ice Accretions,” AIAA–2003–731, January 2003. 

3. Papadakis, M., Gile-Laflin, B.E., “Aerodynamic Performance of a T-tail with Simulated Ice 

Accretions,” January 2000. 

4. Papadakis, M., Alansatan, S., Yeong, H.W., “Aerodynamic Performance of a Swept Wing with Ice 

Accretions,” AIAA–2003–731, January 2003. 

5. Addy, H.E., Broeren, A., Zoeckler, J., Lee, S., “A Wind Tunnel Study of Icing Effects on a Business 

Jet Airfoil,” AIAA–2003–727, January 2003. 

6. Addy, H.E., Chung, J.J., “A Wind Tunnel Study of Icing Effects on a Natural Laminar Flow Airfoil,” 

AIAA–2000–95, January 2000. 

7. Addy, Jr., H.E., “Ice Accretions and Icing Effects for Modern Airfoils,” NASA/TP—2000-210031, 

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

8. Addy, H.E., Potapczuk, M.G., Sheldon, D.W., “Modern Airfoil Ice Accretions,” AIAA–1997–174, 

January 1997. 

9. Busch, G., Broeren, A., Bragg, M., “Aerodynamic Fidelity of Sub-scale Two-Dimensional Ice 

Accretion Simulations,” AIAA–2008–7062, 2008. 

10. Broeren, A.P., Bragg, M.B., “Aerodynamics of Ice Remnants from Protected Surfaces,” AIAA–

2006–0259, January 2006. 

11. Busch, G., Broeren, A., Bragg, M., “Aerodynamic Simulation of a Horn-Ice Accretion on a Subscale 

Model,” AIAA–2007–0087, January 2007. 

12. Broeren, A.P., Addy, H.E., Bragg, M.B., “Effect of Intercycle Ice Accretions on Airfoil 

Performance,” AIAA–2002–0240, January 2002. 

13. Whalen, E.A., Broeren, A.P., Bragg, M.B., “Aerodynamics of Scaled Runback Ice Accretions,” 

Journal of Aircraft, vol. 45, no. 2, March-April 2002, pp. 591–603.Kim, H.S., Bragg, M.B., “Effects 

of Leading-Edge Ice Accretion Geometry on Airfoil Aerodynamics,” AIAA 99-3150, June 1999. 

14. Lee, S., Kim, H.S., Bragg, M.B., “Investigation of Factors that Influence Iced-Airfoil Aerodynamics,” 

AIAA–2000–0099, January 2000. 

15. Lee, S., Dunn, T., Gurbacki, H.M., Bragg, M.B., Loth, E., “An Experimental and Computational 

Investigation of Spanwise-Step-Ice Shapes on Airfoil Aerodynamics,” AIAA No. 98-0490, January 

1998. 

16. Cook, D.E., “Relationships of Ice Shapes and Drag to Icing Condition Dimensionless Parameters,” 

AIAA–2000–0486, January 2000. 

17. Potapczuk, M.G., Berkowitz B.M., “An Experimental Investigation of Multi-Element Airfoil Ice 

Accretion and Resulting Performance Degradation,” AIAA–89–0752, January 1989. 

18. Gonsalez, J.C., Arrington E.A., “Aerodynamic Calibration of the NASA Lewis Icing Research 

Tunnel (1987 Test),” AIAA–98–0633, January 1998. 

19. Wright, W.B., Rutkowski, A., “Validation Results for LEWICE 2.0,” NASA/CR—1999-208690, 

January 1999. 

20. Rae, Jr., R.H., and Pope, A., Low-Speed Wind Tunnel Testing, 2nd ed., John Wiley & Sons, 

New York, 1984, pp. 249. 

21. Anderson, D.N., “Manual of Scaling Methods,” NASA/CR—2004-212875, March 2004. 

22. Shin, J., and Bond, T.H., “Repeatability of Ice Shapes in the NASA Lewis Icing Research Tunnel,” 

Journal of Aircraft, vol. 31, no. 5, pp. 1057–1063. 
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14. ABSTRACT

This paper describes a study performed to investigate how aerodynamic performance degradation progresses with time throughout an 

exposure to icing conditions. It is one of the first documented studies of the effects of ice contamination on aerodynamic performance at 

various points in time throughout an icing encounter. Both a 1.5 and 6 ft chord, two-dimensional, NACA-23012 airfoils were subjected to 

icing conditions in the NASA Icing Research Tunnel for varying lengths of time. At the end of each run, lift, drag, and pitching moment 

measurements were made. Measurements with the 1.5 ft chord model showed that maximum lift and pitching moment degraded more 

rapidly early in the exposure and degraded more slowly as time progressed. Drag for the 1.5 ft chord model degraded more linearly with 

time, although drag for very short exposure durations was slightly higher than expected. Only drag measurements were made with the 6 ft 

chord airfoil. Here, drag for the long exposures was higher than expected. Novel comparison of drag measurements versus an icing scaling 

parameter, accumulation parameter times collection efficiency was used to compare the data from the two different size model. The 

comparisons provided a means of assessing the level of fidelity needed for accurate icing simulation.

15. SUBJECT TERMS

Aircraft icing; Aerodynamic performance; Icing exposure duration 

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