Sam Lee

ASRC Aerospace Corporation, Cleveland, Ohio

Billy Barnhart

Bihrle Applied Research, Inc., Hampton, Virginia

Thomas P. Ratvasky

Glenn Research Center, Cleveland, Ohio

Dynamic Wind-Tunnel Testing of a Sub-Scale Iced 

S-3B Viking

NASA/TM—2012-217684

September 2012

AIAA–2010–7986
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Sam Lee

ASRC Aerospace Corporation, Cleveland, Ohio

Billy Barnhart

Bihrle Applied Research, Inc., Hampton, Virginia

Thomas P. Ratvasky

Glenn Research Center, Cleveland, Ohio

Dynamic Wind-Tunnel Testing of a Sub-Scale Iced 

S-3B Viking

NASA/TM—2012-217684

September 2012

AIAA–2010–7986

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

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Level of Review: This material has been technically reviewed by technical management. 
NASA/TM—2012-217684 1 

Dynamic Wind-Tunnel Testing of a Sub-Scale Iced S-3B Viking 

Sam Lee 

ASRC Aerospace Corporation 

Cleveland, Ohio 44135 

Billy Barnhart 

Bihrle Applied Research, Inc. 

Hampton, Virginia 23666 

Thomas P. Ratvasky1

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

The effect of ice accretion on a 1/12-scale complete aircraft model of S-3B Viking was studied in a 

rotary-balance wind tunnel. Two types of ice accretions were considered: ice protection system failure 

shape and runback shapes that form downstream of the thermal ice protection system. The results showed 

that the ice shapes altered the stall characteristics of the aircraft. The ice shapes also reduced the control 

surface effectiveness, but mostly near the stall angle of attack. There were some discrepancies with the 

data with the flaps deflected that were attributed to the low Reynolds number of the test. Rotational and 

forced-oscillation studies showed that the effects of ice were mostly in the longitudinal forces, and the 

effects on the lateral forces were relatively minor. 

1.0 Introduction 

An S-3B Viking aircraft was considered for possible use at NASA Glenn Research Center for use in 

icing research. This type of aircraft had never been used for icing research, so its flight characteristics in 

an icing environment were unknown. In order to gain insight into the icing effects prior to conducting 

icing research flights, this experimental effort was initiated to determine the aircraft’s iced aerodynamic 

characteristics. 

The effort consisted of 4 parts: 

(1) Determine the ice shapes to test. 

(2) Test the ice shapes on a subscale complete aircraft model in a rotary balance wind tunnel. 

(3) Use the wind tunnel data to develop iced aircraft flight simulation models. 

(4) Refine and validate the flight simulation models through flight testing. 

This method was successfully performed during research efforts for the DHC-6 Twin Otter (Refs. 1 

to 3) and a Cessna business jet (Refs. 4 to 6). 

It was decided to test two ice accretion shapes. The first was the shape that would form on the aircraft 

when the ice protection system has failed (failure shape). The second was the runback shape that forms 

downstream of the leading edge when the aircraft’s thermal ice protection system becomes overwhelmed. 

The failure shape geometry was obtained using LEWICE simulation (Ref. 7). The runback shape geometry 

was simulated with simple aerodynamically equivalent shapes using lessons learned from previous 

research efforts (Refs. 4 and 5). This paper details the testing of the two ice shapes in a rotary balance wind 

tunnel in order to obtain the aerodynamic data needed for the simulation model development. 

 

1

Distinguished Research Associate. 
NASA/TM—2012-217684 2 

Nomenclature 

b aircraft span 

c aircraft chord 

Cl aircraft rolling moment coefficient 

CL aircraft lift coefficient 

CL,max aircraft maximum lift coefficient 

Cm aircraft pitching moment coefficient 

LWC liquid water content 

MVD median volumetric diameter 

OAT outside air temperature 

p aircraft roll rate 

q aircraft pitch rate 

q freestream dynamic pressure 

Re Reynolds number 

V airspeed 

 aircraft angle of attack 

e elevator deflection angle 

f flap deflection angle 

 rotational rate about the velocity vector 

2.0 Experimental Methods 

The test was conducted in the Bihrle Applied Research Large Amplitude Multi-Purpose (LAMP) 

facility in Neuberg a.d. Donau, Germany. It is an open circuit vertical wind tunnel with a 10-ft diameter 

circular test section. It is capable of static, rotational, and forced-oscillation measurements. The forces 

were measured by a six-component strain-gauge balance affixed to the end of a sting and mounted inside 

the aircraft model. The tests were conducted at a dynamic pressure of 2 psf, which corresponded to a 

model Re = 0.10 million. Shown in Figure 1 is the 1/12-scale S-3B model mounted in the test section. The 

model was constructed of fiberglass, balsa, and plywood from a 3D CAD data of the aircraft. The model 

was equipped with moveable control surfaces (aileron, rudder, elevator, flaps, spoilers and speed brakes) 

as well as extended landing gear. A preliminary conceptual model of icing research instrument cluster pod 

was also attached to each wing stores pylon. The ice shapes chosen for testing consisted of the following: 

 Ice protection system failure 

 Runback ice shapes that form downstream of thermal anti-ice system 

The failure shape geometry was generated using LEWICE 3D using the flight conditions shown in 

Table 1. The condition shown below was chosen to represent a typical severe icing encounter during a 

research flight. LEWICE 3D generates ice shapes in series of 2D slices at various spanwise locations. 

These were then lofted to form three-dimensional geometry. Ice shapes geometries were obtained for 

wing and horizontal and vertical stabilizers. These shape were then geometrically scaled from the fullscale to 1/12 scale. Previous experimental studies have shown that airfoils with large leading edge ice 

accretions are insensitive to Reynolds number effects, and simply scaling the ice shapes geometrically for 

subscale testing produces satisfactory results (Refs. 8 and 9). The methodology to develop the failure ice 

 
NASA/TM

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NASA/TM—2012-217684 5 

(a) (b)

Figure 4.—Comparison of airframe manufacturer’s and LAMP lift coefficient data from a previous study. (a) f = 0°.

(b) f = 20°. 

 (deg) 

(a) 

 (deg) 

(b)

Figure 5.—Effect of ice shape on lift. (a) f = 0°. (b) f = 25°. 

 = 14°, with a local maximum lift value of 0.86. It may appear from Figure 5(a) that these ice shapes did 

not significantly degrade the aircraft lift curve. However, these test were conducted at Reynolds number 

of approximately 0.10 million, whereas the full scale aircraft typically operates at 20 million. Thus, in 

order to apply the data to full-scale flight simulation model, the clean lift curve would be expanded to 

give a higher maximum lift and stall angle of attack, as shown in Figure 4. Every iced aerodynamics 

studies have shown that large leading edge iced configurations are Reynolds number insensitive. So, 

when applied to a full-scale aircraft, the differences between the clean and failure iced configurations 

would be much greater. 

Figure 5(b) shows the lift curves with the flaps (both leading and trailing edge) deflected at 25°. As 

expected, deploying the flap resulted in higher lift for both clean and iced configurations. On the clean 

model, it increased the CL,max from 0.90 to 1.38 and decreased the stall angle of attack from 12 to 7°. The 

maximum lift coefficient of the failure shape configuration was increased to 1.5, and the stall angle of 

attack was 12°. These were much higher than the clean model, which was not expected. The runback 

shape configuration had a CL,max of 1.30 with a stall angle of attack of 11°. It was not clear why the failure 

configuration had such a high CL,max when the flap was deployed. 

 

(deg)

CL

LAMP Uncorrected

LAMP Shifted

Flight

10

0.2

 (deg)

CL

LAMP Uncorrected

LAMP Shifted

Flight

10

0.2

0 5 10 15 20

CL

0.0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

1.6

Clean 

Failure 

Runback 

0 5 10 15 20

CL

0.0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

1.6

Clean 

Failure 

Runback 
NASA/TM—2012-217684 6 

 (deg) 

(a)

 (deg) 

(b)

Figure 6.—Effect of ice shape on flap effectiveness. (a)  = 0°. (b)  = 12°. 

The unexpected effect of flap deflection is better illustrated in Figure 6, which shows the effect of ice 

shapes on flap effectiveness. The figure shows the aircraft lift coefficient verses flap deflection angles at 

constant aircraft angles of attack. Figure 6(a) shows the flap effectiveness at  = 0° (which is in the linear 

range of the aircraft lift curve), and Figure 6(b) shows the flap effectiveness at  = 12° (which is at the 

stall angle of attack of the clean aircraft). Figure 6(a) shows that in the linear range of the lift curve, the 

ice shapes had very little effect on the flap effectiveness, with the failure shape configuration showing a 

slightly reduce values, as indicated by its slope. Figure 6(b) shows that at for the clean aircraft, the flap 

effectiveness at stall angle of attack was unchanged from the linear range of the lift curve. However, with 

the failure shape configuration, the flap was three times more effectiveness in increasing lift at  = 12° 

than the clean configuration. For the runback shape configuration, it was 1.5 times more effective. 

 One explanation for this anomaly is that the ice shapes on the wing are helping the boundary layer 

remain attached to the wing when the flaps are deployed. Usually, ice shapes on aircraft wing causes the 

boundary layer to separate from the wing and stall at a lower angle of attack. However, it appears that at 

this low Reynolds number, the opposite may be happening. In the previous research efforts with the Twin 

Otter and the Cessna business jet, this behavior was not observed. However, these aircraft only had 

trailing edge flaps and not leading edge flaps. Because a Reynolds number and geometric scaling studies 

were not performed on the S-3B wing panel (unlike in previous efforts), the exact cause of this behavior 

could not be determined. 

3.2 Pitching Moment 

 Figure 7 shows the effect of the ice shapes on the S-3B Viking model aircraft pitching moment. On the 

clean aircraft, the pitching moment coefficient decreased (i.e., more nose down) linearly with angle attack 

at angles of attack between 0 and 11°. At  = 11°, there was an abrupt nose up pitching moment as the 

aircraft stalled. Both of the ice shape configurations showed more positive pitching moment coefficient 

(when compared to the clean model) at angles of attack between 5° and stall. The sudden nose up pitching 

moment on the failure configuration did not occur until  = 14 and 12° for the failure and runback shapes, 

respectively. These corresponded to breaks in the lift curve slopes on Figure 5(a), indicating the angles of 

attack where the aircraft stalled. The break in the pitching moment coefficient was more gradual with the 

runback configuration than that of the clean aircraft and failure configuration, indicating a softer, more 

gradual stall process. 

0 5 10 15 20 25

CL

0.0

0.2

0.4

0.6

0.8

1.0

Clean 

Failure 

Runback 

0 5 10 15 20 25

CL

0.6

0.8

1.0

1.2

1.4

1.6

Clean 

Failure 

Runback 
NASA/TM—2012-217684 7 

 (deg) 

Figure 7.—Effect of ice shape on pitching moment. 

e (deg) 

(a)

e (deg) 

(b)

Figure 8.—Effect of ice shape on elevator effectiveness.  = 5°. (a) f = 0°. (b) f = 25°.

Figure 8 shows the effect of ice shapes on the elevator effectiveness, or the ability of the elevators to 

change the aircraft pitching moment. Figure 8(a) shows the aircraft pitching moment verses elevator 

deflection angle with the f = 0°, and Figure 8(b) shows the pitching moment verses elevator angle at 

f = 25°. The elevator deflection angles are different in Figures 8(a) and 8(b) because the S-3 has different 

elevator deflection range at different flap settings. Both of the figures are at  = 5°, which was well 

before stall. Figure 8 shows that there was a nearly linear relationship between the aircraft pitching 

moment and the elevator deflection angle. Figure 8(a) shows that ice shapes resulted in a slightly 

decreased elevator effectiveness (indicated by slightly reduced magnitude of Cm–e slope) for f = 0°. At 

f = 25°, all three ice configurations exhibited similar elevator effectiveness. 

Figure 9 shows the effect of ice shapes on the pitch authority. It shows the change in the pitching 

moment the aircraft is able to obtain with full elevator deflection at given lift coefficient. It clearly shows 

that the presence of ice shape shifted the curves to the left, indicating a reduction in the flight envelope. 

0 5 10 15 20

C

m

-0.20

-0.18

-0.16

-0.14

-0.12

-0.10

-0.08

-0.06

-0.04

-0.02

Clean 

Failure 

Runback 

-30 -20 -10 0 10 20

Cm

-0.3

-0.2

-0.1

0.0

0.1

0.2

Clean 

Failure 

Runback 

-30 -20 -10 0 10 20

Cm

-0.4

-0.3

-0.2

-0.1

0.0

0.1

0.2

0.3

Clean 

Failure 

Runback 
NASA/TM—2012-217684 8 

Cm

Figure 9.—Effect of ice shape on pitch authority. Maximum 

elevator deflection. 

3.3 Rolling Moment 

Figure 10 shows the effect of ice shapes on the S-3B Viking model aircraft rolling moment. On the 

clean aircraft, there was a sudden roll off at  = 11°, as it stalled. With the failure shape configuration, the 

roll off was delayed until  = 14° and was not as severe. With the runback configuration, the roll off 

started at  = 12°, and it was even more severe than the clean configuration. 

Figure 11 shows the effect of ice shapes on the aircraft roll authority. It shows the change in the 

rolling moment the aircraft if able to generate with full aileron deflection at a given lift coefficient. On the 

clean aircraft, there was a reduction in the roll authority occurring at CL = 0.35 but it was gradual until 

CL = 0.82, at which there was a rapid decrease in the roll authority. The initial gradual reduction in the roll 

authority may be due to the flow starting to separate over the aileron. The rapid decrease that occurred at 

higher lift values was likely due to the wing stall. Similar behavior was observed with the two iced 

configurations. With the failure shape, the roll authority was similar to that of the clean aircraft at low lift 

values. However, at CL values between 0.7 and 0.9, it exhibited lower roll authority than the clean aircraft. 

However, at CL > 0.90, both clean and failure configuration wing was stalled and showed similar roll 

authority. When the runback shape configuration was installed, the aircraft exhibited reduced roll 

authority even at low lift values and continued up to stall. It is important to keep in mind that for the 

simulation of a full scale aircraft, the clean values would be shifted as discussed above, thus creating an 

even larger decrease in roll authority with the failure and runback shapes compared to the clean aircraft. 

3.4 Rotational Data 

The effects of the icing on the rotational characteristics were mostly observed in the stall region due 

to the differences in their stall characteristics and their effects on the rotational rolling moment 

coefficient. Below stall, there was little or no influence of the ice condition. Figure 12(a) shows that at 

 = 5° (well before stall), the clean and iced aircraft rotational characteristics were nearly identical. The 

differences were observed near stall angles of attack, as shown in Figure 12(b). The differences were 

small and likely due to different stall characteristics. At angle of attack well past stall (Fig. 12(c)) changes 

in the rotational rates did not affect the rolling moment for both clean and iced configurations. The effect 

of ice conditions on rotational yawing moment were observed in the same angle of attack region as roll, 

but were smaller and of less importance. 

 

0.00 0.05 0.10 0.15 0.20 0.25 0.30

CL

0.0

0.2

0.4

0.6

0.8

1.0

1.2

Clean 

Failure 

Runback 
NASA/TM—2012-217684 9 

 (deg) 

Figure 10.—Effect of ice shape on rolling moment. 

Cl

Figure 11.—Effect of ice shape on roll authority. 

(a) 

b/2V (b) 

b/2V

(c)

b/2V

Figure 12.—Effect of ice shape on rotational rolling moment. (a)  = 5°. (b)  = 12°. (c)  = 20°. 

0 5 10 15 20

Cl

-0.025

-0.020

-0.015

-0.010

-0.005

0.000

0.005

Clean 

Failure 

Runback 

0.000 0.005 0.010 0.015 0.020 0.025 0.030 0.035

CL

0.0

0.2

0.4

0.6

0.8

1.0

1.2

Clean 

Failure 

Runback 

-0.3 -0.2 -0.1 0.0 0.1 0.2 0.3

Cl

-0.10

-0.05

0.00

0.05

0.10

Clean 

Failure 

Runback 

-0.3 -0.2 -0.1 0.0 0.1 0.2 0.3

Cl

-0.10

-0.05

0.00

0.05

0.10

Clean 

Failure 

Runback 

-0.3 -0.2 -0.1 0.0 0.1 0.2 0.3

Cl

-0.10

-0.05

0.00

0.05

0.10

Clean 

Failure 

Runback 
NASA/TM—2012-217684 10 

3.5 Forced Oscillation 

Forced-oscillation tests were performed about all three body axis for the aircraft. The effect of ice 

shapes on the forced-oscillation data was generally small and occurred mostly in the stall region, which 

was similar to what was observed in the rotational data. 

Figure 13 shows the effect of roll oscillation on the rolling moment. Figure 13(a) shows that ice 

shapes had very little effect at  = 5°. The only differences observed were near stall angle of attack, as 

shown in Figure 13(b). Like the rotational data, the effects of ice shapes were small. Figure 14 shows the 

effect of pitch oscillation on the pitching moment. The effects were very small for both pre-stall 

(Fig. 14(a)) and stall (Fig. 14(b)) angles of attack. Analysis of other forced-oscillation data showed 

similar trend of ice shapes having little effect when compared to the clean aircraft. 

(a) (b)

Figure 13.—Effect of ice shape on forced-oscillation rolling moment. (a)  = 5°. (b).  = 12°. 

(a) (b) 

Figure 14.—Effect of ice shape on forced-oscillation pitching moment. (a)  = 5°. (b).  = 12°. 

pb2V

-0.10 -0.05 0.00 0.05 0.10

Cl

-0.04

-0.02

0.00

0.02

0.04

0.06

Clean 

Failure 

Runback 

pb2V

-0.10 -0.05 0.00 0.05 0.10

Cl

-0.04

-0.02

0.00

0.02

0.04

0.06

Clean 

Failure 

Runback 

qc2V

-0.010 -0.005 0.000 0.005 0.010

Cm

-0.3

-0.2

-0.1

0.0

0.1

Clean 

Failure 

Runback 

qc2V

-0.010 -0.005 0.000 0.005 0.010

Cm

-0.3

-0.2

-0.1

0.0

0.1

Clean 

Failure 

Runback 
NASA/TM—2012-217684 11 

4.0 Conclusions 

Wind tunnel measurements of an iced 1/12-scale complete aircraft model of an S-3B Viking were 

made in a rotary-balance wind tunnel to obtain data for iced-aircraft flight simulation models. Two 

types of ice accretions were tested: ice protection system failure shape and runback shape that forms 

downstream of the thermal ice-protection system. 

The iced configuration did not severely degrade the aircraft performance when compared to the 

sub-scale clean configuration. This was expected because of the low Reynolds number of this test. The 

clean aircraft at the test Reynolds number would have substantially lower maximum lift coefficient and 

stall angle of attack than at flight Reynolds number. However, the numerous previous tests have shown 

that airfoils with large leading edge ice shapes are relatively unaffected by Reynolds number variation. 

Thus, the iced performance values obtained in the subscale test would be similar to the full-scale 

Reynolds number values. However, this would need to be validated by a full-scale test (i.e., high 

Reynolds number wind tunnel or flight test). 

The data with the flaps deflected at 25° showed the failure ice shape configuration with a much 

higher CL,max than the clean configuration. This was not expected and was speculated to be an artifact of 

the low Reynolds number of this test. Because a Reynolds number scaling study was not done for this 

aircraft model, the exact reason for this could not be determined. A further study, including Reynolds 

number and geometry scaling effects, is required if the flap-deflected data is to be incorporated into a 

flight simulation model. 

The ice shapes had some measurable effects on control surface effectiveness, but mostly near stall. 

The effects of ice shapes on the rotational characteristics were not significant at pre-stall angles of 

attack. Most significant differences were observed in the stall regions in the rolling moment. The most 

significant effect of ice on the forced oscillation was measured in the rolling-moment coefficient due to 

roll rate near the stall angles of attack. The ice configurations had little effect on pitch and yaw 

damping values. 

References 

1. Ratvasky, T.P., Ranaudo, R.J., Barnhart, B.P., Dickes, E.G., and Gingras, D.R., “Development 

and Utility of a Piloted Flight Simulator for Icing Effects Training,” 41st AIAA Aerospace 

Sciences Meeting and Exhibit, Jan. 6–9, 2003, Reno, NV, AIAA Paper 2003–0022, 

NASA/TM—2003-212116. 

2. Gingras, D.R., Dickes, E.G., Ratvasky, T.P., and Barnhart, B.P., “Modeling of In-Flight Icing 

Effects for Pilot Training,” AIAA Modeling and Simulation Technologies Conference and Exhibit, 

Aug. 5–8, 2002, Monterey, CA, AIAA Paper 2002-4605. 

3. Papadakis, M., Gile-Laflin, B.E., Youssef, G.M., and Ratvasky, T.P., “Aerodynamic Scaling 

Experiments With Simulated Ice Accretions,” 39th AIAA Aerospace Sciences Meeting and Exhibits, 

Jan. 8–11, 2001, Reno, NV, AIAA Paper 2001–0833. 

4. Lee, S., Barnhart, B.P., Ratvasky, T.P., Dickes, E.G., and Thacker, M., “Dynamic Wind Tunnel 

Testing of a Sub-Scale Iced Business Jet,” 44th AIAA Aerospace Sciences Meeting and Exhibit, 

Jan. 9–12, 2006, Reno, NV, AIAA Paper 2006–0261, NASA/TM—2006-214268. 

5. Lee, S., Ratvasky, T.P., Thacker, M., and Barnhart, B.P., “Geometry and Reynolds Number Scaling 

on an Iced Business Jet Wing,” 43rd AIAA Aerospace Sciences Meeting and Exhibits, Jan. 10–13, 

2005, Reno, NV, AIAA Paper 2005–1066, NASA/TM—2005-1066. 

6. Ratvasky, T.P., Barnhart, B.P., Lee, S., and Cooper, J., “Flight Testing an Iced Business Jet for 

Flight Simulation Model Validation,” 45th AIAA Aerospace Sciences Meeting and Exhibit, 

Jan. 8–11, 2007, Reno, NV, AIAA Paper 2007–0089, NASA/TM—2007-214936. 

7. Bidwell, C.S., “Icing Analysis of the NASA S3 Icing Research Aircraft Using LEWICE3D Version 

2,” SAE Paper 07ICE137, Sept. 2007. 
NASA/TM—2012-217684 12 

8. Addy, H.E., Broeren, A.P., Zoeckler, J.G., and Lee, S., “A Wind Tunnel Study of Icing Effects on a 

Business Jet Airfoil,” 41st AIAA Aerospace Science Meeting and Exhibit, Jan. 6–9, 2003, Reno, 

NV, AIAA Paper 2003–0727. 

9. Broeren, A.P., Bragg, M.B., Addy, H.E., Lee, S., Moens, F., and Guffond, D., “Effect of HighFidelity Ice Accretion Simulation on the Performance of a Full-Scale Airfoil Model,” 46th AIAA 

Aerospace Sciences Meeting and Exhibit, Jan. 7–10, 2008, Reno, NV, AIAA Paper 2008–0434. 

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Dynamic Wind-Tunnel Testing of a Sub-Scale Iced S-3B Viking 

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14. ABSTRACT

The effect of ice accretion on a 1/12-scale complete aircraft model of S-3B Viking was studied in a rotary-balance wind tunnel. Two types 

of ice accretions were considered: ice protection system failure shape and runback shapes that form downstream of the thermal ice 

protection system. The results showed that the ice shapes altered the stall characteristics of the aircraft. The ice shapes also reduced the 

control surface effectiveness, but mostly near the stall angle of attack. There were some discrepancies with the data with the flaps deflected 

that were attributed to the low Reynolds number of the test. Rotational and forced-oscillation studies showed that the effects of ice were 

mostly in the longitudinal forces, and the effects on the lateral forces were relatively minor. 

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Icing; Aerodynamics 

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