Andy P. Broeren and Harold E. Addy, Jr.

Glenn Research Center, Cleveland, Ohio

Sam Lee

Vantage Partners, LLC, Brook Park, Ohio

Marianne C. Monastero

University of Illinois, Urbana-Champaign, Urbana, Illinois

Validation of 3-D Ice Accretion Measurement 

Methodology for Experimental 

Aerodynamic Simulation

NASA/TM—2015-218724

July 2015

AIAA–2014–2614
NASA STI Program . . . in Profile

Since its founding, NASA has been dedicated 

to the advancement of aeronautics and space science. 

The NASA Scientific and Technical Information (STI) 

Program plays a key part in helping NASA maintain 

this important role.

The NASA STI Program operates under the auspices 

of the Agency Chief Information Officer. It collects, 

organizes, provides for archiving, and disseminates 

NASA’s STI. The NASA STI Program provides access 

to the NASA Technical Report Server—Registered 

(NTRS Reg) and NASA Technical Report Server—

Public (NTRS) thus providing one of the largest 

collections of aeronautical and space science STI in 

the world. Results are published in both non-NASA 

channels and by NASA in the NASA STI Report 

Series, which includes the following report types:

• TECHNICAL PUBLICATION. Reports of 

completed research or a major significant phase 

of research that present the results of NASA 

programs and include extensive data or theoretical 

analysis. Includes compilations of significant 

scientific and technical data and information 

deemed to be of continuing reference value. 

NASA counter-part of peer-reviewed formal 

professional papers, but has less stringent 

limitations on manuscript length and extent of 

graphic presentations.

• TECHNICAL MEMORANDUM. Scientific 

and technical findings that are preliminary or of 

specialized interest, e.g., “quick-release” reports, 

working papers, and bibliographies that contain 

minimal annotation. Does not contain extensive 

analysis.

• CONTRACTOR REPORT. Scientific and 

technical findings by NASA-sponsored 

contractors and grantees.

• CONFERENCE PUBLICATION. Collected 

papers from scientific and technical 

conferences, symposia, seminars, or other 

meetings sponsored or co-sponsored by NASA.

• SPECIAL PUBLICATION. Scientific, 

technical, or historical information from 

NASA programs, projects, and missions, often 

concerned with subjects having substantial 

public interest.

• TECHNICAL TRANSLATION. Englishlanguage translations of foreign scientific and 

technical material pertinent to NASA’s mission.

For more information about the NASA STI 

program, see the following:

• Access the NASA STI program home page at 

http://www.sti.nasa.gov

• E-mail your question to help@sti.nasa.gov

• Fax your question to the NASA STI 

Information Desk at 757-864-6500

• Telephone the NASA STI Information Desk at

757-864-9658

• Write to:

NASA STI Program

Mail Stop 148

NASA Langley Research Center

Hampton, VA 23681-2199
Andy P. Broeren and Harold E. Addy, Jr.

Glenn Research Center, Cleveland, Ohio

Sam Lee

Vantage Partners, LLC, Brook Park, Ohio

Marianne C. Monastero

University of Illinois, Urbana-Champaign, Urbana, Illinois

Validation of 3-D Ice Accretion Measurement 

Methodology for Experimental 

Aerodynamic Simulation

NASA/TM—2015-218724

July 2015

AIAA–2014–2614

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

6th Atmospheric and Space Environments Conference

sponsored by the American Institute of Aeronautics and Astronautics

Atlanta, Georgia, June 16–20, 2014
Acknowledgments

The Atmospheric Environment Safety Technologies Project of the NASA Aviation Safety Program with Ron Colantonio as project 

manager provided support for this research. The authors thank the crew of the NASA Icing Research Tunnel for their support of 

the ice-accretion testing and molding portion of this work. Lloyd Utt provided invaluable assistance with post-processing of the 

laser-scan data and development of the rapid-prototype artificial ice shapes under the NASA Undergraduate Student Researcher’s 

Program. The Federal Aviation Administration provided support for ice-casting fabrication and wind-tunnel testing at University 

of Illinois. The authors gratefully acknowledge Jeff Diebold, Brian Woodard, Dave Foley and Mike Bragg at University of Illinois 

for their contributions to this research.

Available from

Level of Review: This material has been technically reviewed by technical management. 

NASA STI Program

Mail Stop 148

NASA Langley Research Center

Hampton, VA 23681-2199

National Technical Information Service

5285 Port Royal Road

Springfield, VA 22161

703-605-6000

This report is available in electronic form at http://www.sti.nasa.gov/ and http://ntrs.nasa.gov/
NASA/TM—2015-218724 1 

Validation of 3-D Ice Accretion Measurement Methodology for 

Experimental Aerodynamic Simulation

Andy P. Broeren and Harold E. Addy, Jr. 

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135 

Sam Lee

Vantage Partners, LLC

Brook Park, Ohio 44135 

Marianne C. Monastero

University of Illinois, Urbana-Champaign 

Urbana, Illinois 61801 

Abstract

Determining the adverse aerodynamic effects due to ice accretion often relies on dry-air wind-tunnel 

testing of artificial, or simulated, ice shapes. Recent developments in ice-accretion documentation 

methods have yielded a laser-scanning capability that can measure highly three-dimensional (3-D) 

features of ice accreted in icing wind tunnels. The objective of this paper was to evaluate the aerodynamic 

accuracy of ice-accretion simulations generated from laser-scan data. Ice-accretion tests were conducted 

in the NASA Icing Research Tunnel using an 18-in. chord, two-dimensional (2-D) straight wing with 

NACA 23012 airfoil section. For six ice-accretion cases, a 3-D laser scan was performed to document the 

ice geometry prior to the molding process. Aerodynamic performance testing was conducted at the 

University of Illinois low-speed wind tunnel at a Reynolds number of 1.8×106

 and a Mach number of 0.18 

with an 18-in. chord NACA 23012 airfoil model that was designed to accommodate the artificial ice 

shapes. The ice-accretion molds were used to fabricate one set of artificial ice shapes from polyurethane 

castings. The laser-scan data were used to fabricate another set of artificial ice shapes using rapid 

prototype manufacturing such as stereolithography. The iced-airfoil results with both sets of artificial ice 

shapes were compared to evaluate the aerodynamic simulation accuracy of the laser-scan data. For five of 

the six ice-accretion cases, there was excellent agreement in the iced-airfoil aerodynamic performance 

between the casting and laser-scan based simulations. For example, typical differences in iced-airfoil 

maximum lift coefficient were less than 3 percent with corresponding differences in stall angle of 

approximately 1° or less. The aerodynamic simulation accuracy reported in this paper has demonstrated 

the combined accuracy of the laser-scan and rapid-prototype manufacturing approach to simulating ice 

accretion for a NACA 23012 airfoil. For several of the ice-accretion cases tested, the aerodynamics is 

known to depend upon the small, three-dimensional features of the ice. These data show that the laserscan and rapid-prototype manufacturing approach is capable of replicating these ice features within the 

reported accuracies of the laser-scan measurement and rapid-prototyping method; thus providing a new 

capability for high-fidelity ice-accretion documentation and artificial ice-shape fabrication for icing 

research.
NASA/TM—2015-218724 2 

Nomenclature

c airfoil chord length

Cd drag coefficient

Cd,casting drag coefficient for the three-dimensional casting 

Cd,rpm drag coefficient for the RPM simulation

ΔCd,rms root-mean-square percent difference in drag coefficient between the RPM simulation and the 

three-dimensional casting simulation over a given angle of attack range

Cl lift coefficient

Cl,max maximum lift coefficient, coincident with stalling angle

Cm quarter-chord pitching moment 

Cp pressure coefficient

k ice roughness height or ice thickness

M freestream Mach number

N number of angles of attack 

Re freestream Reynolds number based on chord 

α angle of attack

αstall stalling angle of attack, coincident with Cl,max

LWC liquid water content

MVD median volumetric diameter 

RPM rapid prototype manufacturing 

SLA stereolithography 

1.0 Introduction

Ice accretion and its aerodynamic effect on airfoil and wing performance are important to the 

continued design, certification and safe operation of aircraft in the icing environment. Icing wind tunnels 

are a critical engineering tool for simulating natural ice accretion on aircraft surfaces. As described by 

Lee et al. (Ref. 1), the chief product of such facilities is the ice accretion that forms on various test articles 

ranging from airplane wings, rotor blades, engine inlets, radomes and other flight hardware. Bosetti et al.

(Ref. 2) discuss many important aspects of documenting ice-shape characterization. They point out that 

ice shapes are often used by other groups for follow-on computational or experimental aerodynamic 

studies; development of design criteria or requirements; engineering tool development, improvement or 

validation. Therefore, documentation of the resulting ice accretion is a key piece of data in icing-windtunnel tests. In a companion paper, Lee et al. (Ref. 3) describe continued progress on the development of 

a 3-D ice-accretion measurement methodology using laser-scanning systems, including a geometric 

evaluation of its accuracy. This paper focuses on an aerodynamic approach for evaluating the accuracy of 

ice accretion measured using the laser-scanning approach on a subscale straight wing. 

Lee et al. (Ref. 1) describe the documentation of ice-accretion geometries in icing-wind-tunnel tests 

using many different techniques. These techniques include standard (qualitative) photography, 

quantitative photography, tracings, and various molding and casting techniques using impression foam, 

wax, plaster and other materials. The best current technology for capturing three-dimensional features of 

ice accretion is the mold and casting method. This has been used for many years, but was improved at 

NASA Glenn (then Lewis) during the 1980s using more robust materials to improve accuracy and 

durability (Ref. 4). The main disadvantages to the mold and casting method are that it can be time 

consuming both during the icing test and in post-test production. Cost can be significant in some cases but 

generally scales with the size of the ice accretion to be documented. There is no electronic or digitized 
NASA/TM—2015-218724 3 

record of the ice-accretion geometry, only a physical casting. Additional measurements are required in 

order to generate quantitative information about the ice-accretion geometry. Lee et al. (Ref. 3) report on 

recent developments with laser-scanning based methods that are capable of directly measuring and 

digitally archiving the 3-D features of ice-accretion geometry. 

One motivation for ice-accretion documentation is to evaluate the resulting aerodynamic effect.

According to Broeren et al. (Ref. 5) the most common way to acquire iced-airfoil and wing data is to use 

geometric representations of ice accretion in a dry-air wind tunnel or in flight. These geometric 

representations are often referred to as “artificial ice shapes” or “ice-accretion simulations.” The highest 

fidelity ice-accretion simulations are castings developed from the molding process described in the 

previous paragraph. These ice-accretion castings maintain the major features of the ice, including the 

detailed surface roughness and the spanwise and chordwise variations. Typically, these castings are 

attached to wings and airfoils, and instrumented to obtain high-fidelity aerodynamic data (Refs. 6 to 11). 

This is an expensive process and not practical for many situations, but does generate benchmark data for 

iced-airfoil and wing research. Broeren et al. (Ref. 5) describe a recently completed comprehensive study 

of aerodynamic simulation of ice accretion that utilized the mold and casting method as benchmark data 

in order to evaluate the aerodynamic accuracy of lower-fidelity ice-accretion simulations. For some cases 

such as initial roughness, spanwise ridges and streamwise ice shapes, it was found that the fine geometric 

details can be important to the resulting aerodynamics. Therefore, it is important to understand the 

accuracy to which 3-D ice-accretion geometry can be measured using new tools such as laser scanning 

described by Lee et al. (Ref. 3).

The objective of the current study is to evaluate the aerodynamic accuracy of ice-accretion 

simulations generated from laser-scan data. Ice-accretion tests were conducted in the NASA Icing 

Research Tunnel (IRT) using an 18-in. chord, 2-D straight wing with NACA 23012 airfoil section. 

Several different types of ice accretions were generated. For the selected cases, a 3-D laser scan was 

performed to document the ice-accretion geometry prior to the molding process. Artificial ice shapes were 

fabricated for aerodynamic testing in the University of Illinois low-speed wind-tunnel. The artificial ice 

shapes were attached to the leading edge of an 18-in. chord NACA 23012 airfoil model. The ice-accretion 

molds were used to fabricate one set of artificial ice shapes from polyurethane castings. The laser-scan 

data were used to fabricate artificial ice shapes using rapid-prototype manufacturing (RPM) such as 

stereolithography. Aerodynamic performance testing was conducted at a Reynolds number of 1.8×106

 and 

a Mach number of 0.18. The iced-airfoil results with both sets of artificial ice shapes were compared to 

evaluate the aerodynamic simulation accuracy of the laser-scan data. It should be noted that the 

aerodynamic comparisons amount to a combined evaluation of not only the laser-scanning measurement 

method, but also the accuracy of the RPM used to generate the artificial ice shapes. For the purposes of 

this paper, uncertainties in the scan data and manufacturing tolerances in the RPM are combined into the 

resulting artificial ice shape. The companion paper by Lee et al. (Ref. 3) contains more information on the 

uncertainty in the scan data alone. 

2.0 Experimental Methods and Apparatus

2.1 Aerodynamic Experiments

All aerodynamic testing was performed using the low-speed, low-turbulence wind tunnel at the 

University of Illinois with the experimental apparatus described by Broeren et al. (Ref. 5) The wind 

tunnel has a 33.6-in. (0.85-m) by 48-in. (1.2-m) test section capable of speeds up to Mach 0.20. An 18-in. 

(0.46-m) chord NACA 23012 airfoil model was designed with interchangeable leading edges that 

accommodated the various ice simulations. There was a baseline leading edge having the NACA 23012 
NASA/TM—2015-218724 4 

profile that was used to document the un-iced, or clean, airfoil performance, and there were two ice 

leading edges with truncated nose geometry that allowed for the attachment of the ice simulations. The 

artificial ice shapes bolted onto the ice leading edges and thus had a rigid, repeatable mounting system. 

Figure 1 shows the NACA 23012 model installed in the University of Illinois wind tunnel. The model had 

a main chordwise row of pressure taps, a secondary chordwise row, and a set of spanwise taps on the 

upper surface. Both sets of artificial ice shapes (castings and RPM shapes) were also instrumented with 

static pressure taps.

The model lift and pitching-moment data were acquired from a force balance and by integration of 

airfoil surface static pressures measured by an electronically scanned pressure system. Good agreement 

between these methods was generally obtained. Locating and installing pressure orifices in and around 

3-D ice features sometimes led to spurious Cp points on the artificial ice shape at certain angles of attack.

These spurious data could affect the lift and pitching-moment coefficients determined from integrated 

surface pressure data. For this reason, and for consistency with the results presented in Broeren et al.

(Ref. 5) the lift and pitching moment data for the clean configuration were obtained from the surface 

pressures, whereas the data reported for the iced configurations were obtained from the force balance in 

this paper. Momentum-deficit methods were used to compute the drag coefficient from total-pressure 

measurements collected by a traversable wake rake. This wake rake is shown downstream of the model in 

Figure 1. Surface-oil flow visualization was also conducted for selected cases. A light coat of mineral oil 

treated with fluorescent dye was applied to the surface of the model in the region of interest. The tunnel 

flow condition was then set and maintained for a period of time sufficient to allow the local wall shear 

stress to cause the oil to flow on the surface. The resulting oil-flow patterns indicated boundary-layer 

transition and regions of attached and separated flow. Monastero and Bragg (Refs. 12 and 13) describe 

pressure-sensitive paint (PSP) measurements that were also conducted for some of the artificial ice-shape 

configurations. 

Figure 1.—NACA 23012 airfoil model installed in the University of Illinois windtunnel test section with artificial ice shape installed on leading edge.
NASA/TM—2015-218724 5 

TABLE 1.—ESTIMATED EXPERIMENTAL UNCERTAINTIES 

Aerodynamic 

quantity

Reference 

value

Absolute 

uncertainty

Relative uncertainty, 

% 

α 4.16° ±0.02° ±0.48

Cl Balance 0.548 ±0.00019 ±0.35

Cm Balance –0.0020 ±0.00023 ±12.1

Cp –0.962 ±0.0045 ±0.47

Cd Wake 0.0071 ±0.00014 ±1.9

Busch (Ref. 14), Monastero (Ref. 12), and Broeren et al. (Ref. 5) also provide more details about the 

experimental setup, including the data uncertainty analysis. The experimental uncertainties were 

calculated for 20:1 odds using the methods of Kline and McClintock (Ref. 15) and Coleman and Steele 

(Ref. 16). Sample uncertainties are provided in Table 1 for the clean-model configuration. These 

uncertainties were acceptable for this investigation. The relative uncertainty in the pitching-moment 

coefficient appears to be large because the reference value is small. The methods of Allen and Vincenti 

(Ref. 17) and Barlow, Rae and Pope (Ref. 18) were used to correct the angle of attack, lift coefficient, 

pitching-moment coefficient and drag coefficient for wind-tunnel wall boundary effects. All data were 

collected at a Reynolds number of Re = 1.8×106

 and a corresponding Mach number of M = 0.18. 

2.2 Ice-Accretion Experiments

Ice-accretion tests were conducted in the NASA Icing Research Tunnel (IRT) using an 18-in. chord, 

2-D straight wing with NACA 23012 airfoil section. Several different types of ice accretions were 

generated, and a subset of these were selected for aerodynamic evaluation. Table 2 provides the 

aerodynamic and icing cloud conditions for the selected cases. The first column, “Ice-Shape 

Classification” refers to the aerodynamic classification described by Bragg et al. (Ref. 19). For this 

investigation, there were two roughness shapes, two streamwise ice shapes, one spanwise-ridge ice shape 

and one horn ice shape. For these ice shapes listed in Table 2, a 3-D laser scan was performed to 

document the ice-accretion geometry prior to the molding process. The ice-accretion molds were used to 

fabricate one set of artificial ice shapes from polyurethane castings. Monastero (Ref. 12) provides a 

detailed description of the mold-and-casting process used in this research. The laser-scan data were used 

to fabricate the RPM artificial ice shapes. Lee et al. (Ref. 3) provide a detailed description of process used 

to develop the RPM artificial ice shapes from the laser-scan data. Care was taken to ensure that the 

castings and RPM shapes corresponded to identical spanwise locations on the IRT model. The spanwise 

length of the finished artificial ice shapes was limited to one-third of the aerodynamic model span, so that 

three identical sections were required to cover the span of the model in the Illinois wind tunnel. The 

artificial ice shapes had pressure taps installed at equivalent locations for the castings and RPM shapes.

Figure 2 shows pictures of the instrumented segments of the ED1977 streamwise ice shape simulations 

for comparison. 

In all cases the RPM shapes were manufactured using stereolithography (SLA). Data provided by the 

manufacturer indicated an accuracy of ±0.015 in. with a minimum layer thickness of 0.005 in. The 

manufacturer also stated that feature sizes less than 0.025 in. in size were not guaranteed to build with 

SLA. In the case of the ED1983 shape, an additional method, High-Density PolyJet was used. The 

objective here was to evaluate any difference in the methods used for rapid prototyping. Since the 

ED1983 shape had the smallest roughness features, this was determined to be the best candidate for this 

evaluation. Data provided by the manufacturer indicated an accuracy of ±0.005 in. with a minimum layer 

thickness of 0.0006 in. The minimum feature size for High-Density PolyJet was 0.012 in. 
NASA/TM—2015-218724 6 

TABLE 2.—SUMMARY OF IRT TEST CONDITIONS FOR

ICE SHAPES SELECTED FOR AERODYNAMIC EVALUATION

Ice-shape

classification

IRT run

number

Airspeed,

kt

α MVD,

μm

LWC,

g/m3

Total 

temperature,

°F/°C

Static 

temperature,

°F/°C

Spray time,

min

Roughness 1 ED1974 200 2.0° 15 0.75 28.0/–2.2 18.5/–7.5 0.5

Roughness 2 ED1983 200 2.0° 30 0.40 0.0/–17.8 –9.5/–23.1 1.0

Streamwise 1 ED1966 175 5.0° 15 0.30 0.0/–17.8 –7.3/–21.8 5.0

Streamwise 2 ED1977 200 2.0° 30 0.40 0.0/–17.8 –9.5/–23.1 5.0

Spanwise ridge ED1967 175 1.0° 15 0.64 24.0/–4.4 16.8/–8.4 9.5

Horn ED1978 200 2.0° 15 0.75 28.0/–2.2 18.5/–7.5 5.0

As indicated in Table 2, a total of six different ice accretions were selected for aerodynamic testing 

following the aerodynamic classifications described by Bragg et al. (Ref. 19). Graphical information for 

each case is shown in Figure 3 to Figure 8. The section cuts shown in the figures were extracted from the 

laser-scan data for each configuration. In the case of the artificial ice shapes, the laser scan was acquired 

with the artificial ice shape bolted to the leading edge of the NACA 23012 wind-tunnel model. Therefore, 

the laser-scan data accurately represent the geometry as it was tested in the aerodynamic wind tunnel.

Also shown for reference is a section cut extracted from the original ice accretion scan. Each figure shows 

a 3-D rendering of the artificial ice shape solid model used for production of the RPM shapes. Finally, a 

photograph of the ice accretion is also included. 

Two cases of ice roughness were selected to evaluate the accuracy of the laser-scan and RPM 

approach to developing artificial ice shapes. The ED1974 case was a glaze-ice roughness while ED1983 

was a rime-ice roughness. The latter had very small roughness feather features virtually invisible in 

Figure 4 that were thought to be challenging to the laser-scanning and RPM processes. For both cases, the 

ice thickness was very small and, as a result, it is difficult to meaningfully compare the section cuts in the 

regions dominated by roughness The ice was thicker (and much smoother) in the region of the stagnation 

point. The section cuts in Figure 3 for the ED1974 case compare favorably in this region. In contrast, the 

section cuts in Figure 4 for the ED1983 case indicate that there was a significant decrease in stagnation 

point ice thickness for the casting configuration. This discrepancy was attributed to misalignment of the 

ice mold and leading-edge geometry during the casting process. Lee et al. (Ref. 3) also note that the cast 

shapes were often smaller than the RPM shapes and the original ice accretion thus indicating shrinkage 

during the mold and casting process. For the ED1983 case, there is a more favorable comparison among 

both RPM shapes with the ice accretion. More discussion of the roughness features is included with 

presentation of the aerodynamic data in the next section.

Two different streamwise ice shape cases were selected to evaluate the accuracy of the laser-scan and 

RPM approach. The ED1966 case shown in Figure 5 was characterized by a main ice shape located near 

the airfoil leading edge with rime feathers downstream. These rime feather features were small, but still 

visible in the Figure 5 images. The ED1977 case shown in Figure 6 was characterized by a main ice shape 

that was more conformal to the airfoil leading edge with much larger rime feather features. The larger 

rime feathers also extended farther downstream, both effects being attributable to the larger drop size

(MVD = 30 μm) used to form the ice accretion. The section cuts shown in Figure 5 and Figure 6 

demonstrate good agreement among the artificial ice-shape configurations with the original ice accretion.
NASA/TM—2015-218724 7 

A spanwise-ridge ice shape was also selected to evaluate the accuracy of the laser-scan and RPM 

approach. The ED1967 case shown in Figure 7 was formed with an electrothermal surface heater mounted 

to the model leading edge. Icing conditions and heater settings were adjusted to create conditions where 

water flowed downstream over the heater on both the upper and lower surfaces of the model. The 

objective was to simulate the development of a spanwise-ridge ice accretion that may result from a 

thermal ice-protection system. The images in Figure 7 illustrate the spanwise ridges that formed on both 

upper and lower surfaces and that there was good agreement among the ice-shape configurations with the 

ice accretion itself. The electrothermal surface heater did affect the laser-scan data and resulting artificial 

ice shapes which Lee et al. (Ref. 3) explain in their paper.

Finally, a horn ice shape was selected to evaluate the accuracy of the laser-scan and RPM approach.

The ED1978 case shown in Figure 8 is a typical glaze ice accretion with both upper- and lower-surface 

horns. The section cuts do indicate some differences in the upper-surface horn geometry for the casting 

configuration relative to the ice-accretion data and RPM configurations. The upper-surface horn of the 

casting was slightly shorter and more rounded. Also, there was a significant amount of ice missing from 

the casting near the base of the upper-surface horn at x/c ≈ 0.0, y/c ≈ 0.025. In the case of hand tracing, 

such differences might simply be attributable to spanwise variations in the ice accretion, illustrated in the 

accompanying images. This explanation is unsatisfactory in this case, since extracting section cuts from 

the laser-scan data is much more precise than hand tracing. As Lee et al. (Ref. 3) describes, the 

measurements indicate that were was some shrinkage and warpage in the casting that led to the geometric 

differences shown in Figure 8. Lee et al. (Ref. 3) provides a more detailed description and analysis of the 

geometric differences between the laser-scan and RPM shapes and the corresponding casting 

configurations. 

Figure 2.—Pressure tap segments of ED1977 streamwise ice-shape simulations; left—casting, right—RPM artificial 

ice shape.
NASA/TM—2015-218724 8 

Figure 3.—ED1974 roughness ice shape: section cuts extracted from laser-scan data (left); rendering of 

RPM artificial ice shape (top right) and ice-accretion photograph (bottom right).

Figure 4.—ED1983 roughness ice shape: section cuts extracted from laser-scan data (left); rendering of 

RPM artificial ice shape (top right) and ice-accretion photograph (bottom right).

x/c

y/c

-0.02 0.00 0.02 0.04 0.06 0.08 0.10

-0.04

-0.02

0.00

0.02

0.04

0.06

NACA 23012

ED1974 Ice Accretion

ED1974 Casting

ED1974 RPM

x/c

y/c

-0.02 0.00 0.02 0.04 0.06 0.08 0.10

-0.04

-0.02

0.00

0.02

0.04

0.06

NACA 23012

ED1983 Ice Accretion

ED1983 Casting

ED1983 RPM-SLA

ED1983 RPM-PolyJet HD
NASA/TM—2015-218724 9 

Figure 5.—ED1966 streamwise ice shape: section cuts extracted from laser-scan data (left); rendering of 

RPM artificial ice shape (top right) and ice-accretion photograph (bottom right).

Figure 6.—ED1977 streamwise ice shape: section cuts extracted from laser-scan data (left); rendering of 

RPM artificial ice shape (top right) and ice-accretion photograph (bottom right).

x/c

y/c

-0.04 0.00 0.04 0.08 0.12 0.16

-0.06

-0.04

-0.02

0.00

0.02

0.04

0.06

0.08

0.10

NACA 23012

ED1966 Ice Accretion

ED1966 Casting

ED1966 RPM

x/c

y/c

-0.04 0.00 0.04 0.08 0.12 0.16

-0.06

-0.04

-0.02

0.00

0.02

0.04

0.06

0.08

0.10

NACA 23012

ED1977 Ice Accretion

ED1977 Casting

ED1977 RPM
NASA/TM—2015-218724 10

Figure 7.—ED1967 spanwise-ridge ice shape: section cuts extracted from laser-scan data (left); rendering 

of RPM artificial ice shape (top right) and ice-accretion photograph (bottom right).

Figure 8.—ED1978 horn-ice shape: section cuts extracted from laser-scan data (left); rendering of RPM 

artificial ice shape (top right) and ice-accretion photograph (bottom right).

x/c

y/c

0.00 0.04 0.08 0.12 0.16 0.20 0.24

-0.08

-0.04

0.00

0.04

0.08

0.12

NACA 23012

ED1967 Ice Accretion

ED1967 Casting

ED1967 RPM

x/c

y/c

-0.04 -0.02 0.00 0.02 0.04 0.06 0.08

-0.04

-0.02

0.00

0.02

0.04

0.06

NACA 23012

ED1978 Ice Accretion

ED1978 Casting

ED1978 RPM
NASA/TM—2015-218724 11

3.0 Results

The main objective of this research was to evaluate the combined accuracy of the laser scanning/rapid 

prototyping process for iced-aerodynamic simulation. The effect of the ice shape on airfoil performance is 

compared for both a casting simulation and the corresponding RPM simulation generated from the laserscan data. As described in the Introduction, the ice casting has been considered the best simulation of the 

actual ice accretion for dry-air aerodynamic testing. Therefore, the RPM simulation is compared against 

the casting simulation as the accepted standard. These comparisons are shown for each of the ice shapes 

described in the previous section after a brief presentation of the clean airfoil results.

3.1 Clean Airfoil

A series of clean-airfoil experiments were performed to ensure that the experimental apparatus 

yielded results consistent with Broeren et al. (Ref. 5) and Busch (Ref. 14). The NACA 23012 airfoil lift, 

pitching-moment and drag coefficient variation with angle of attack is shown in Figure 9 for both the 

present data and previous data from Busch (Ref. 14). There was excellent agreement in the lift coefficient 

and pitching-moment coefficients based upon the integrated surface pressure. In Figure 9, the maximum 

lift coefficient was 1.49 for the present data compared to 1.48 for Busch (Ref. 14) The stalling angle of 

attack was 14.4° for both data sets. Individual surface pressure distributions at identical angles of attack 

were also compared and demonstrated equally good agreement. There were larger differences in the drag 

coefficient between the two data sets. In general, the present data had higher drag coefficient. The reason 

for this was not investigated in detail. The differences were attributed to changes in the model surface 

condition and alignment of the removable leading edges. Since the clean airfoil drag is highly dependent 

upon the surface condition, any deterioration in the surface condition, or small misalignments in the 

removable leading edge could result in increased drag. Furthermore, any simulated ice shape would have 

significantly larger flow disturbance thus minimizing any issues with the clean model surface condition. 

Therefore, the disagreement in the clean airfoil drag did not compromise the objectives of the present 

experiments.

Figure 9.—Clean NACA 23013 airfoil aerodynamic performance comparison for the present data with past 

results from Busch (Ref. 14) at Re = 1.8×106 and M = 0.18.

α (deg.)

-12 -8 -4 0 4 8 12 16

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

1.6

-0.08

-0.06

-0.04

-0.02

0.00

0.02

0.04

0.06

0.08

0.10

0.12

0.14

0.16

Cl Cm

α (deg.)

-10 -8 -6 -4 -2 0 2 4 6 8 10 12 14 16

0.000

0.005

0.010

0.015

0.020

0.025

0.030

Busch Data

Present Data

Cd
NASA/TM—2015-218724 12

3.2 Ice Roughness 

The aerodynamic results presented in Figure 10 show the NACA 23012 lift, pitching-moment and 

drag coefficients plotted against angle of attack for the clean airfoil and for the ED1974 roughness 

simulated with the casting and the RPM shape. This relatively small amount of roughness had a large 

effect on the airfoil performance. For the casting simulation, Cl,max and αstall were reduced to 1.04° and 

10.3°, respectively. Overall, the agreement in aerodynamic performance between the casting and the RPM 

simulations is considered to be very good. The RPM simulation had slightly lower Cl,max = 1.01 at the 

same stalling angle. This lower Cl,max value was preceded by a small decrease in the lift-curve slope and 

accompanying higher drag beginning at approximately α = 7° 

To facilitate comparisons in drag coefficient, Busch et al. (Ref. 20) developed a parameter to compare 

Cd over an appropriate angle of attack range. The angle of attack range was that over which Cl varied 

approximately linearly with α. The parameter, ΔCd,rms, is presented as a percentage and is computed by 

determining the root-mean-square of the percent difference between the casting Cd and the corresponding 

RPM simulation Cd at each angle of attack in the predefined linear range (a total of N angles of attack):

N

x

C

CC

C

N

i

i

castingd

i

castingd

i

rpmd

d rms

∑

= 













 −

=∆ 1

2

,

,,

,

%100

For the ED1974 case, ΔCd,rms was 3.8 percent calculated on the interval –4° ≤ α ≤ 8°. This value is 

very low indicating excellent agreement between the drag coefficient of the casting simulation and the 

RPM simulation. A summary of the key performance-based parameters is provided in Table 3 for 

comparison. Previous reports (Refs. 5, 20 to 23) have tabulated these parameters for artificial ice shapes 

using different simulation methods having varying degrees of geometric fidelity representing the ice 

accretion. The results reported in Table 3 for the ED1974 RPM configuration are very good compared to 

these previous results, further supporting the excellent aerodynamic comparison between the casting and 

RPM simulations.

Figure 10.—Aerodynamic performance comparison for the ED1974 roughness with casting and RPM simulations on 

the NACA 23012 airfoil at Re = 1.8×106 and M = 0.18.

α (deg.)

-12 -8 -4 0 4 8 12 16

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

1.6

-0.08

-0.06

-0.04

-0.02

0.00

0.02

0.04

0.06

0.08

0.10

0.12

0.14

0.16

Cl Cm

α (deg.)

-10 -8 -6 -4 -2 0 2 4 6 8 10 12 14 16

0.00

0.01

0.02

0.03

0.04

0.05

0.06

Clean

ED1974 Casting

ED1974 RPM

Cd
NASA/TM—2015-218724 13

TABLE 3.—SUMMARY OF AERODYNAMIC ACCURACY FOR THE RPM SIMULATIONS

Ice-shape configuration RPM Cl,max – Casting Cl,max RPM αstall – Casting αstall ΔCd,rms,

%

α-range for 

ΔCd,rms

ED1974 RPM –0.026 (–2.5%) –0.01° 3.8 –4° ≤ α ≤ 8°

ED1983 RPM-SLA 0.071 (6.3%) 1.03° 11.1 –4° ≤ α ≤ 8°

ED1983 RPM-PolyJet HD 0.126 (11.2%) 1.04° 12.5 –4° ≤ α ≤ 8°

ED1966 RPM –0.001 (–0.1%) 0.00° 2.9 –4° ≤ α ≤ 8°

ED1977 RPM 0.002 (0.2%) 0.00° 2.5 –4° ≤ α ≤ 8°

ED1967 RPM –0.042 (–2.9%) –1.01° 5.0 –4° ≤ α ≤ 8° 

ED1978 RPM –0.034 (–4.8%) –0.03° 10.0 –4° ≤ α ≤ 6°

ED1978 RPM of Casting 0.011 (1.6%) 0.00° 1.8 –4° ≤ α ≤ 6°

Figure 11.—Pressure distribution comparison for the ED1974 roughness with casting and RPM simulations on 

the NACA 23012 airfoil at α = 9.3°, Re = 1.8×106 and M = 0.18.

While the overall performance results shown in Figure 10 and key parameters summarized in Table 3 

represent two measures of the aerodynamic simulation accuracy, it is also important to consider flowfield 

details. In this investigation, flowfield differences were assessed through comparison of the surface 

pressure distributions and surface-oil flow visualization. An example of the former is shown in Figure 11 

for each of the configurations at α = 9.3°. The comparison between the airfoil with the casting and the 

airfoil with the RPM shape is very good, with Cp being virtually identical, except in the leading-edge 

region. Differences here were attributable to differences in the individual roughness element geometry 

relative to the location of the pressure orifice. Compared to the clean airfoil Cp, the data show flowfield 

behavior that is typical for ice roughness. The simulated ice roughness led to a reduction in the suction 

pressure on the upper surface from the leading edge downstream to x/c ≈ 0.20. There was also a 

divergence of Cp downstream of x/c ≈ 0.85 that is consistent with boundary-layer separation near the 

airfoil trailing edge. The presence of this trailing-edge separation was confirmed in the surface-oil flow 

visualization results. Taken together, these results demonstrate the capability of the laser scanning to 

x/c

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

-5.0

-4.0

-3.0

-2.0

-1.0

0.0

1.0

Clean

ED1974 Casting

ED1974 RPM

C

p
NASA/TM—2015-218724 14

accurately measure the small roughness features for the ED1974 ice accretion, as well as the ability of the 

stereolithography RPM to accurately reproduce those same roughness features on the artificial ice shape. 

There are practical limitations in terms of the roughness sizes than can be measured and reproduced 

in an RPM artificial ice shape. The ED1983 case was selected in order to assess these practical 

limitations. This ice accretion was the most challenging case owing to the very small-scale roughness 

associated with the icing condition. The lift, pitching-moment and drag coefficient data plotted in 

Figure 12 show more significant differences among the casting, stereolithography and PolyJet HD 

simulations than was previously observed for the ED1974 case. For the airfoil with the ED1983 casting 

simulation, Cl,max and αstall were reduced from the clean airfoil case to 1.12° and 11.3°, respectively. As 

noted in Table 3, the SLA simulation resulted in a Cl,max that was 6.3 percent higher than the casting while 

Cl,max for the PolyJet HD simulation was 11.2 percent higher with a corresponding increase in stall angle 

of approximately 1° for both configurations. There were also significant differences in drag coefficient 

over a large angle of attack range. These differences are quantified in terms of the ΔCd,rms parameter 

evaluated on the interval –4° ≤ α ≤ 8° in Table 3. The ΔCd,rms values of 11.1 and 12.5 percent for the SLA 

and PolyJet HD shapes were significantly larger than the 3.8 percent reported for the ED1974 case. 

The distribution of surface pressure is plotted in Figure 13 for each of the ED1983 configurations at 

α = 9.3°. The plot shows that the largest differences in Cp between the casting configuration and the RPM 

configurations were observed in the upper-surface suction pressures from the leading edge downstream to 

x/c ≈ 0.10. In this region, Cp for the RPM configurations more closely matched that of the clean airfoil 

than for the casting configuration. This provides further evidence that the small roughness associated with 

the ice accretion was not accurately simulated by the laser scan and RPM approach for this ice accretion 

case.

Figure 12.—Aerodynamic performance comparison for the ED1983 roughness with casting and RPM simulations on 

the NACA 23012 airfoil at Re = 1.8×106 and M = 0.18.

α (deg.)

-12 -8 -4 0 4 8 12 16

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

1.6

-0.08

-0.06

-0.04

-0.02

0.00

0.02

0.04

0.06

0.08

0.10

0.12

0.14

0.16

Cl Cm

α (deg.)

-10 -8 -6 -4 -2 0 2 4 6 8 10 12 14 16

0.00

0.01

0.02

0.03

0.04

0.05

0.06

Clean

ED1983 Casting

ED1983 RPM-SLA

ED1983 RPM-PolyJet HD

Cd
NASA/TM—2015-218724 15

Figure 13.—Pressure distribution comparison for the ED1983 roughness with casting and RPM simulations on the 

NACA 23012 airfoil at α = 9.3°, Re = 1.8×106 and M = 0.18.

Quantifying the size and distribution of small ice roughness is very difficult. Figure 14 shows images 

taken from the laser-scan surfaces for the ED1974 and ED1983 ice accretion. While qualitative, these 

images do show the differences in the roughness characteristics between these two cases. For the ED1974 

case, the typical roughness sizes estimated form the scan data were on the order of 0.010 to 0.015 in.

(0.25 to 0.38 mm). This size is within both the stated accuracy of the laser-scanner measurement system

and stereolithography prototyping method. On the other hand, the typical roughness sizes for the ED1983 

case estimated from the scan data were an order of magnitude smaller, in the range of 0.001 to 0.0015 in. 

(0.025 to 0.038 mm). This size range is at the accuracy limit of the laser-scanning measurement system 

and below the accuracy of either the stereolithography or PolyJet HD prototyping methods. In future 

work, the roughness size and distribution should be better quantified in order to further document the 

limitations of the laser scan and RPM approach. As noted by McClain and Kreeger (Ref. 24), image 

analysis techniques have been used in the past to determine the dimensions of roughness elements from 

photographic images. These methods are laborious and not practical to process a large number of cases.

McClain and Kreeger (Ref. 24) demonstrated a method for determining roughness size characteristics 

utilizing the high-resolution laser-scan data of the type described in this paper. This method could be 

applied to these data in order to better characterize the roughness size distribution and better quantify the 

limit of the laser scan/RPM approach.

x/c

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

-4.0

-3.0

-2.0

-1.0

0.0

1.0

Clean

ED1983 Casting

ED1983 RPM-SLA

ED1983 RPM-PolyJet HD

C

p
NASA/TM—2015-218724 16

ED1974 Ice-Accretion Photograph ED1983 Ice-Accretion Photograph

ED1974 Laser-Scan Surface Data Image ED1983 Laser-Scan Surface Data Image

ED1974 Casting Photograph ED1983 Casting Photograph

Figure 14.—Images of surface roughness for the ED1974 (left) and ED1983 (right) ice accretions.

3.3 Streamwise Ice

Two streamwise ice cases were evaluated in this investigation. Aerodynamic results for the ED1966 

case are shown in Figure 15. The stalling characteristics were nearly identical for the airfoil with the 

casting simulation or with the RPM simulation. Both artificial ice shapes resulted in Cl,max = 1.00 at 

αstall = 10.3°. It is interesting to note that these values are very similar to the effect of the ED1974 

roughness shape on the airfoil stalling characteristics. Aerodynamic results for the ED1977 case are 

shown in Figure 16. Once again, the stalling characteristics were nearly identical for the airfoil with the 

casting simulation or with the RPM simulation. Both artificial ice shapes resulted in Cl,max = 1.04 at 

αstall = 11.3°. For this case, the stalling angle was 1° higher than for the ED1966 case. Comparisons in the 

iced-airfoil drag between the casting and RPM shapes were also very good for both streamwise ice cases.

This comparison is reflected in the very low values of the ΔCd,rms parameter summarized in Table 3. Not 

surprisingly, a review of the surface pressure data and flow visualization results also demonstrated 

excellent matching. 
NASA/TM—2015-218724 17

Figure 15.—Aerodynamic performance comparison for the ED1966 streamwise ice-shape casting and RPM 

simulations on the NACA 23012 airfoil at Re = 1.8×106 and M = 0.18.

Figure 16.—Aerodynamic performance comparison for the ED1977 streamwise ice-shape casting and RPM 

simulations on the NACA 23012 airfoil at Re = 1.8×106 and M = 0.18.

The aerodynamic simulation of streamwise ice was investigated in detail by Broeren et al. (Ref. 5) 

and Busch et al. (Refs. 14, 20, and 23). This type of aerodynamic matching requires accurate reproduction 

of certain characteristics of the roughness and feather features of the ice accretion. The reproduction of 

the gross ice shape is actually less important for streamwise ice where the main ice accretion tends to be 

an extension of the airfoil leading edge. The detailed 3-D geometric comparison data (Ref. 3) indicated 

that there was some shrinkage associated with the casting simulation as well as the loss of certain 

roughness and feather features relative to the original ice accretion. Since the reproduction of the gross (or 

main) ice shape is less aerodynamically important for streamwise ice, the apparent shrinkage had little to 

α (deg.)

-12 -8 -4 0 4 8 12 16

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

1.6

-0.08

-0.06

-0.04

-0.02

0.00

0.02

0.04

0.06

0.08

0.10

0.12

0.14

0.16

Cl Cm

α (deg.)

-10 -8 -6 -4 -2 0 2 4 6 8 10 12 14 16

0.00

0.01

0.02

0.03

0.04

0.05

0.06

Clean

ED1966 Casting

ED1966 RPM

Cd

α (deg.)

-12 -8 -4 0 4 8 12 16

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

1.6

-0.08

-0.06

-0.04

-0.02

0.00

0.02

0.04

0.06

0.08

0.10

0.12

0.14

0.16

Cl Cm

α (deg.)

-10 -8 -6 -4 -2 0 2 4 6 8 10 12 14 16

0.00

0.01

0.02

0.03

0.04

0.05

0.06

Clean

ED1977 Casting

ED1977 RPM

Cd
NASA/TM—2015-218724 18

no effect on the aerodynamic results. Even though geometric differences were also observed in the 

roughness and feather regions, these differences were not aerodynamically significant. Therefore, these 

results further demonstrate the ability of the laser scanning/rapid prototyping method to accurately 

reproduce the roughness and feather features for this type of ice accretion for the purposes of 

aerodynamic performance evaluation.

3.4 Spanwise-Ridge Ice

The aerodynamic performance results for the ED1967 spanwise-ridge ice are shown in Figure 17. The 

main adverse effect in aerodynamic performance for this ice shape is the large increase in drag coefficient.

The maximum lift coefficient was 1.45 for the casting simulation compared to 1.48 for the clean 

configuration. The stalling angle for the airfoil with the casting simulation was increased by approximately 

2° to 16.4°. This type of aerodynamic behavior with short spanwise ridge ice was documented in detail by 

Whalen et al. (Refs. 25 to 28) and Broeren et al. (Refs. 29 to 31). The spanwise-ridge geometry in this case

acts to delay stall relative to the clean airfoil at this Reynolds number. However, at increased Reynolds 

number, the clean airfoil maximum lift coefficient and stall angle would both increase, while the iced-airfoil 

performance would be expected to remain approximately the same (Ref. 31). For example, at 

Re = 16.0 × 106 and M = 0.20, the clean NACA 23012 has Cl,max = 1.82 and αstall = 18.1° (Refs. 5 and 31) At 

this Reynolds number, the spanwise-ridge ice accretion would have a significant effect on the airfoil 

performance (as would the other ice accretion cases previously described). 

The spanwise-ridge geometry in this case is similar to runback ice accretion that is observed on 

airfoils and wings equipped with thermal ice protection systems that provide continuous heat input to the 

leading edge. Runback ice accretion occurs when the amount of heat being supplied to the leading edge is 

less than the amount required to evaporate all of the impinging water. The water that does not evaporate 

“runs back” downstream and freezes, typically on an unheated portion of the wing. Broeren et al. (Ref. 5). 

summarize the significant aerodynamic factors. Because the ice ridge is located well downstream of the 

airfoil leading edge, a boundary layer develops on the smooth leading-edge region upstream of the ridge.

The characteristics of this boundary layer then depend upon many factors including the airfoil geometry,

Figure 17.—Aerodynamic performance comparison for the ED1967 spanwise-ridge ice-shape casting and RPM 

simulations on the NACA 23012 airfoil at Re = 1.8×106 and M = 0.18.

α (deg.)

-12 -8 -4 0 4 8 12 16

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

1.6

-0.08

-0.06

-0.04

-0.02

0.00

0.02

0.04

0.06

0.08

0.10

0.12

0.14

0.16

Cl Cm

α (deg.)

-10 -8 -6 -4 -2 0 2 4 6 8 10 12 14 16

0.00

0.01

0.02

0.03

0.04

0.05

0.06

Clean

ED1967 Casting

ED1967 RPM

Cd
NASA/TM—2015-218724 19

pressure distribution and Reynolds number. For airfoils like the NACA 23012 with high leading-edge 

suction pressure, an ice ridge located farther downstream would be located in a region where there would 

be a significant adverse pressure gradient on the clean airfoil. The resulting implication for aerodynamic 

simulation is that the artificial ice shapes must replicate much of the fine details of the runback ice 

accretion geometry (Ref. 31). Therefore, the good agreement in aerodynamic performance between the 

casting and RPM simulations shown in Figure 17 implies that there was also good agreement in the 

geometry of these two simulations. This favorable comparison is also reflected in the performance metrics 

listed in Table 3. References 5, 29 to 31 provide additional listings of these parameters for various 

spanwise-ridge ice simulations further supporting this evaluation.

Comparisons of the surface pressure distributions and surface-oil flow visualizations were also 

performed to gauge the simulation accuracy. The pressure distribution comparison in Figure 18 for 

α = 6.2°. clearly shows the near-identical match between the casting and RPM shape on the airfoil. The 

pressure distribution is typical for a short spanwise ridge where the main effect is immediately in the 

vicinity of the ridge itself. On the upper surface, there was a local deceleration of the flow at the upstream 

extent of the ridge (x/c ≈ 0.11), followed by the flow acceleration over the top of the ridge corresponding 

to the region of minimum pressure (x/c ≈ 0.13). Within a short distance downstream of the ridge 

(x/c ≈ 0.20), the pressure distribution was nearly the same as on the clean airfoil. Similar characteristics 

were also observed for the lower-surface ridge. This indicated that the aerodynamic effect of the ridge 

was confined locally to its x/c position. The spanwise ridge did not generate large flow separation that 

would have significantly altered the pressure distribution. In fact, the most significant difference between 

the casting and RPM simulations occurred on the upper surface, upstream of the ridge. These differences 

were most likely attributable to the contour accuracy and surface finish of the leading-edge geometry in 

this region (i.e., x/c < 0.10). As noted in Lee et al. (Ref. 3), the contour accuracy in this region was 

affected by the electrothermal surface heater used to generate the runback ice ridge. Given the factors 

described in the previous paragraph, such differences could affect the resulting flow over the ridge and 

subsequent performance. This may explain why the casting simulation had a stall angle 1° higher than for 

the RPM simulation. 

Figure 18.—Pressure distribution comparison for the ED1967 spanwise-ridge ice shape with casting and RPM 

simulations on the NACA 23012 airfoil at α = 6.2°., Re = 1.8×106 and M = 0.18.

x/c

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

-4.0

-3.0

-2.0

-1.0

0.0

1.0

Clean

ED1967 Casting

ED1967 RPM

C

p
NASA/TM—2015-218724 20

3.5 Horn Ice

The aerodynamic performance effects of the horn-ice simulation on the NACA 23012 airfoil are 

shown in Figure 19. This case had the largest adverse effect on the airfoil performance of all cases in this 

investigation. The maximum lift coefficient for the airfoil with the casting was 0.70 at the stalling angle 

of 7.2°. For this ice shape the stall angle was taken at 7.2° based upon the surface pressure distribution 

that showed a significant amount of lift being generated on the forward portion of the airfoil, despite the 

break in Cm at α = 6.2°. This is also consistent with the determination of Cl,max and αstall used in Broeren et 

al. (Ref. 5). Based upon an initial glance at Figure 19, it may appear that there was good agreement 

between the performance results for the airfoil with the casting simulation versus the RPM simulation.

However, a closer review of the data showed that this was not the case. 

The first indication of the poor comparison was based upon an evaluation of the performance-based 

metrics listed in Table 3. These values were compared to the values reported in Broeren et al. (Ref. 5) for 

similar horn ice accretion. In that report, the artificial ice shapes that were compared to the casting 

simulation were a 2-D smooth shape and a simple-geometry shape. The 2-D smooth shape was a 

spanwise extrusion of one section cut of the ice accretion while the simple-geometry simulation only had 

rectangular protuberances (or “spoilers” ) representing the maximum height and angle of the upper and 

lower-surface ice horns. Both artificial ice shapes were tested with and without grit roughness applied to 

the surface, thus representing four different configurations. The resulting performance results were 

compared to the airfoil with the casting simulation. The variation in Cl,max ranged from 0.2 percent for the 

2-D smooth simulation to –3.0 percent for the simple-geometry simulation with added roughness. This 

range is significantly smaller than the –4.8 percent difference reported in Table 3 for the present study. 

Since the RPM shape should accurately represent all aerodynamically relevant 3-D features of the ice 

accretion, a much closer match in Cl,max was expected. Likewise for the ΔCd,rms parameter calculated over 

the same angle of attack interval (–4.0° ≤ α ≤ 8.0°), the range from Broeren et al. (Ref. 5) was 9.4 to 

11.9 percent. While the present value of 10 percent falls within this range, a smaller value was expected.

Looking across the range within Table 3, ΔCd,rms values higher than 10 percent were only observed for the 

ED1983 which has already been described as a poor comparison. 

Figure 19.—Aerodynamic performance comparison for the ED1978 horn ice-shape casting and RPM simulations 

on the NACA 23012 airfoil at Re = 1.8×106 and M = 0.18.

α (deg.)

-12 -8 -4 0 4 8 12 16

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

1.6

-0.08

-0.06

-0.04

-0.02

0.00

0.02

0.04

0.06

0.08

0.10

0.12

0.14

0.16

Cl Cm

α (deg.)

-10 -8 -6 -4 -2 0 2 4 6 8 10 12 14 16

0.00

0.02

0.04

0.06

0.08

0.10

Clean

ED1978 Casting

ED1978 RPM

Cd
NASA/TM—2015-218724 21

In addition to these performance metrics, the poor comparison for the ED1978 horn ice case was also 

observed in the surface pressure distributions and flow visualization. An example of the former is shown 

in Figure 20 for α = 6.2°. The pressure distributions for both the casting simulation and the RPM 

simulation show the distinctive “plateau” region on the upper surface between the leading edge and 

x/c ≈ 0.08. This region of approximately constant pressure corresponds to the separated flow region aft of 

the upper-surface horn. The key difference in the pressure distributions occurred in the pressure recovery 

region on the airfoil upper surface. The casting simulation had a more rapid recovery indicating a smaller, 

or shorter, separated flow region relative to the RPM shape. This observation was corroborated by the 

surface oil flow visualization images in Figure 21. These images show the mean chordwise separation 

bubble reattachment location for each configuration (yellow line). There was some spanwise variation in 

this location because of the three-dimensional geometry of the artificial ice shapes. For the casting 

configuration the mean reattachment location was at x/c ≈ 0.20 compared to x/c ≈ 0.25 for the RPM 

configuration. The pressure tap row had a spanwise location between 16- and 18-in. from the tunnel floor. 

At this spanwise location, the mean reattachment location for the casting configuration was still at 

x/c ≈ 0.20 compared to x/c ≈ 0.30 for the RPM configuration. The flow visualization results confirm that 

the separated flow region for the casting configuration was smaller, or shorter, than for the RPM 

configuration. 

All of the aerodynamic data presented are consistent with the geometry comparison shown in Figure 8 

where the upper-surface horn of the casting simulation was shorter and more rounded than for the RPM 

simulation. Both of these factors are known from previous research (Refs. 14, 32, and 33) to lead to 

higher Cl,max and lower Cd due to a decrease in the size of the separation bubble. As described in Section 

2.0, Lee et al. (Ref. 3) suggest that there were anomalies associated with the casting that led to the 

geometric differences shown in Figure 8. These geometry data indicate that the RPM simulation was a 

better overall match to the ice-accretion geometry. 

Figure 20.—Pressure distribution comparison for the ED1978 horn ice shape with casting and RPM simulations 

on the NACA 23012 airfoil at α = 6.2°, Re = 1.8×106 and M = 0.18.

x/c

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

-3.0

-2.0

-1.0

0.0

1.0

Clean

ED1978 Casting

ED1978 RPM

C

p
NASA/TM—2015-218724 22

Figure 21.—Surface-oil flow visualization comparison for the ED1978 horn ice shape with casting (left) and 

RPM (right) simulations on the NACA 23012 airfoil at α = 6.2°, Re = 1.8×106 and M = 0.18. Yellow line 

denotes the separation bubble mean reattachment location. Horizontal scales mark the x/c location and 

vertical scales mark height from the tunnel flow in inches. Figure adapted from Monastero (Ref. 12). 

Owing to this discrepancy, the aerodynamic evaluation of the horn-ice simulation was investigated

further. Unfortunately, it was not possible to fabricate another casting simulation from the original mold 

of the ice accretion. This mold no longer exists since it had to be rent asunder in the process of removing 

the casting from the mold. An alternative approach was utilized whereby an RPM simulation based upon 

the laser-scan data of the casting was built. Aerodynamic testing was then conducted in follow-on 

campaign. The results are summarized in Figure 22 comparing the performance of the airfoil with the 

ED1978 casting simulation (from Figure 19) along with the RPM simulation based upon the laser-scan of 

the casting. Comparing Figure 19 and Figure 22, it is clear that the RPM shape based upon the casting 

geometry demonstrated much better aerodynamic comparison. This observation is quantified in Table 3 

that shows differences in Cl,max, αstall and ΔCd,rms that are now well within the range of the other cases, 

(excluding the ED1983 configuration). Further evidence of the close matching is provided in Figure 23 

surface pressure distributions where there is nearly an exact match for this angle of attack. The only minor 

differences appear to be at pressure orifice locations on the ice shape itself that can be influenced by local 

three-dimensional features. These results clearly demonstrate the ability of the laser scanning and RPM 

processes to accurately reproduce horn-ice aerodynamics.
NASA/TM—2015-218724 23

Figure 22.—Aerodynamic performance comparison for the ED1978 horn ice-shape casting simulation and RPM 

simulation based upon the casting geometry on the NACA 23012 airfoil at Re = 1.8×106 and M = 0.18.

Figure 23.—Pressure distribution comparison for the ED1978 horn ice-shape casting simulation and RPM simulation

based upon the casting geometry on the NACA 23012 airfoil at α = 6.2°, Re = 1.8×106 and M = 0.18.

4.0 Conclusion

Determining the adverse aerodynamic effects due to ice accretion continues to rely on dry-air windtunnel testing of artificial, or simulated, ice shapes. Recent developments in ice-accretion documentation 

methods have yielded a laser-scanning capability that can capture highly three-dimensional (3-D) features 

of experimental ice accretion in icing wind-tunnel tests. The three-dimensional laser-scan data can 

α (deg.)

-12 -8 -4 0 4 8 12 16

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

1.6

-0.08

-0.06

-0.04

-0.02

0.00

0.02

0.04

0.06

0.08

0.10

0.12

0.14

0.16

Cl Cm

α (deg.)

-10 -8 -6 -4 -2 0 2 4 6 8 10 12 14 16

0.00

0.02

0.04

0.06

0.08

0.10

Clean

ED1978 Casting

ED1978 RPM of Casting

Cd

x/c

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

-3.0

-2.0

-1.0

0.0

1.0

Clean

ED1978 Casting

ED1978 RPM of Casting

C

p
NASA/TM—2015-218724 24

subsequently be used to build fully three-dimensional RPM shapes of the ice accretion. The objective of 

this paper was to evaluate the aerodynamic accuracy of RPM ice-accretion simulations generated from 

laser-scan data. Ice-accretion tests were conducted in the NASA IRT using an 18-in. chord, 2-D straight 

wing with NACA 23012 airfoil section. For the selected ice-accretion cases, a 3-D laser scan was 

performed to document the ice geometry prior to the molding process. The ice-accretion molds were used 

to fabricate one set of artificial ice shapes from polyurethane castings. The laser-scan data were used to 

fabricate another set of artificial ice shapes using rapid-prototype manufacturing such as 

stereolithography. Aerodynamic performance testing was conducted on an 18-in. chord NACA 23012 

airfoil model at a Reynolds number of 1.8×106

 and a Mach number of 0.18. The iced-airfoil results with 

both sets of artificial ice shapes were compared to evaluate the aerodynamic simulation accuracy of the 

laser-scan data. Consistent with past research, the ice-casting simulation was taken as the accepted 

standard, or benchmark, to which the RPM shapes were compared. This aerodynamic comparison was a 

combined evaluation of not only the laser-scanning measurement method, but also the accuracy of the 

rapid-prototype method used to generate the artificial ice shapes. For the purposes of this paper, 

uncertainties in the scan data and manufacturing tolerances in the rapid-prototyping method are combined 

into the resulting artificial ice shape. 

The aerodynamic comparisons were performed for two cases of ice roughness (ED1974 and 

ED1983), two streamwise ice shapes (ED1966 and ED1977), one spanwise-ridge ice shape (ED1967) and 

one horn-ice shape (ED1978). For five of these six cases, there was excellent agreement in the iced-airfoil 

aerodynamic performance between the casting and RPM simulations. For example, typical differences in 

iced-airfoil maximum lift coefficient were less than 3 percent with corresponding differences in stall 

angle of approximately 1° or less. The glaze-ice roughness ice accretion (ED1974) had estimated 

roughness heights of 0.010 to 0.015 in. (0.25 to 0.38 mm). This size is within both the stated accuracy of 

the laser-scanner measurement system and stereolithography prototyping method. The resulting 

aerodynamic comparison was very good. One case that exhibited poor comparison was the small rime-ice 

roughness ice accretion (ED1983). Typical roughness heights in this case were estimated to be in the 

range of 0.001 to 0.0015 in. (0.025 to 0.038 mm). This size range is at the accuracy limit of the laserscanning measurement system and below the capability of the selected rapid-prototype manufacturing 

methods. 

The aerodynamic simulation accuracy reported in this paper has demonstrated the combined accuracy 

of the laser-scan and RPM approach to simulating ice accretion for a NACA 23012 airfoil. For several of 

the ice-accretion cases tested, the aerodynamics is known to depend upon the small, three-dimensional 

features of the ice. These data show that the laser-scan and RPM approach is capable of replicating these 

ice features within the reported accuracies of the measurement and prototype method; thus providing a 

new capability for high-fidelity ice-accretion documentation and artificial ice-shape fabrication for icing 

research.

References

1. Lee, S., Broeren, A.P., Addy, H.E., Jr., Sills, R., and Pifer, E.M., “Development of 3-D Ice Accretion 

Measurement Method,” AIAA Paper 2012-2938, June 2012; also NASA/TM—2012-217702, Sept. 

2012. 

2. Bosetti, C., Paul, B.P., Jr., Malone, A.M., “Ice Shape Characterization to Aid in Replicating Ice 

Shapes for Subsequent Analysis,” AIAA Paper 2010-7534. 

3. Lee, S., Broeren, A.P., Kreeger, R.E., Potapczuk, M.G., and Utt, L., “Implementation and Validation 

of 3-D Ice Accretion Measurement Methodology,” AIAA paper 2014-2613, June 2014. 
NASA/TM—2015-218724 25

4. Reehorst, A.L., and Richter, G.P., “New Methods and Materials for Molding and Casting Ice 

Formations,” NASA TM-100126, Sept. 1987. 

5. Broeren, A.P., Addy, H.E., Jr., Bragg, M.B., Busch, G.T., Guffond, D., and Montreuil, E., 

“Aerodynamic Simulation of Ice Accretion on Airfoils,” NASA/TP—2001-216929, June 2011.

6. Broeren, A.P., Bragg, M.B., and Addy, H.E., Jr., “Effect of Intercycle Ice Accretions on 

Aerodynamic Performance,” Journal of Aircraft, Vol. 41, No. 1, Jan.-Feb. 2004, pp. 165-174. 

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

a Business Jet Airfoil,” AIAA Paper 2003-0727, Jan. 2003; also NASA/TM—2003-212124, Feb. 

2003. 

8. Papadakis, M., Yeong, H.W., Wong, S.C., Vargas, M., and Potapczuk, M.G., “Aerodynamic 

Performance of a Swept Wing with Ice Accretions,” AIAA Paper 2003-0731, Jan. 2003.

9. Gile-Laflin, B.E., and Papdakis, M., “Experimental Investigation of Simulated Ice Accretions on a 

Natural Laminar Flow Airfoil,” AIAA Paper 2001-0088, Jan. 2001.

10. Gregorek, G., Dresse, J.J., and LaNoe, K., “Additional Testing of the DHC-6 Twin Otter Tailplane 

Iced Airfoil Section in the Ohio State University 7x10 Low Speed Wind Tunnel,” NASA/CR—2000-

209921/Vol. 2, Sept. 2000. 

11. Addy, H.E., Jr., and Chung, J.J., “A Wind Tunnel Study of Icing Effects on a Natural Laminar Flow 

Airfoil,” AIAA Paper 2000-0095, Jan. 2000; also NASA/TM—2000-209775, Jan. 2000.

12. Monastero, M.C., “Validation of 3-D Ice Accretion Documentation and Replication Method 

Including Pressure-Sensitive Paint,” M.S. Thesis, Dept. of Aerospace Engineering, Univ. of Illinois at 

Urbana-Champaign, 2013. 

13. Monastero, M.C., and Bragg, M.B., “Validation of 3-D Ice Accretion Measurement Methodology 

Using Pressure-Sensitive Paint (PSP),” AIAA 6th Atmospheric and Space Environments Conference, 

Atlanta, GA, June 16-20, 2014 (submitted for publication). 

14. Busch, G.T., “Ice Accretion Aerodynamic Simulation on a Subscale Model” M.S. Thesis, Dept. of 

Aerospace Engineering, Univ. of Illinois at Urbana-Champaign, 2006. 

15. Kline, S., and McClintock, F.A, “Describing Uncertainties in Single Sample Experiments,” 

Mechanical Engineering, Vol. 75, No. 1, 1953, pp. 3-8. 

16. Coleman, H.W., and Steele, W.G., Experimentation and Uncertainty Analysis for Engineers, WileyInterscience, New York, 1989, pp.40-118. 

17. 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 782, 1944. 

18. Barlow, J.B., Rae, W.H., Jr., and Pope, A., Low-Speed Wind Tunnel Testing, John Wiley & Sons, 

1999, pp. 328-366. 

19. Bragg, M.B., Broeren, A.P., and Blumenthal, L.A., “Iced-Airfoil Aerodynamics,” Progress in 

Aerospace Sciences, Vol. 41, No. 5, July 2005, pp. 323-418. 

20. Busch, G.T., Broeren, A.P., and Bragg, M.B., “Aerodynamic Fidelity of Sub-scale Two-Dimensional 

Ice Accretion Simulation,” AIAA Paper 2008-7062, Aug. 2008. 

21. Broeren, A.P., Whalen, E.A., Busch, G.T., and Bragg, M.B., “Aerodynamic Simulation of Runback 

Ice Accretion,” Journal of Aircraft, Vol. 47, No. 3, May-June 2010, pp. 924-939, also AIAA Paper 

2009-4261, June 2009. 

22. Broeren, A.P., Whalen, E.A., Busch, G.T., and Bragg, M.B., “Aerodynamic Simulation of Runback 

Ice Accretion,” U.S. Dept. of Transportation/Federal Aviation Administration Rept. DOT/FAA/AR09/26, Dec. 2009, also NASA/TM—2010-215676, Mar. 2010. 

23. Busch, G.T., “Experimental Study of Full-Scale Iced-Airfoil Aerodynamic Performance using 

Subscale Simulations,” Ph.D. dissertation, Dept. of Aerospace Eng., Univ. of Illinois, Urbana, 

Illinois, 2010.

24. McClain, S.T., and Kreeger, R.E., “Assessment of Ice Shape Roughness Using a Self-Organizing 

Map Approach,” AIAA Paper 2013-2546, June 2013. 
NASA/TM—2015-218724 26

25. Whalen, E.A., Broeren, A.P., and Bragg, M.B., “Considerations for Aerodynamic Testing of Scaled 

Runback Ice Accretions,” AIAA Paper 2006-0260, 44th Aerospace Sciences Meeting and Exhibit, 

Reno, NV, Jan. 2006.

26. Whalen, E.A, Broeren, A.P., and Bragg, M.B., “Characteristics of Runback Ice Accretion and their 

Aerodynamic Effects,” U.S. Dept. of Transportation/Federal Aviation Administration Rept. 

DOT/FAA/AR-07/16, April 2007. 

27. Whalen, E.A., “Aerodynamics of Runback Ice Accretions,” Ph.D. dissertation, Dept. of Aerospace 

Eng., Univ. of Illinois, Urbana, Illinois, 2007. 

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

Journal of Aircraft, Vol. 45, No. 3, May-June 2008, pp. 1076-1088. 

29. Broeren, A.P., Whalen, E.A., Busch, G.T., and Bragg, M.B., “Aerodynamic Simulation of Runback 

Ice Accretion,” AIAA Paper 2009-4261, June 2009.

30. Broeren, A.P., Whalen, E.A., Busch, G.T., and Bragg, M.B., “Aerodynamic Simulation of Runback 

Ice Accretion,” DOT/FAA/AR-09/26, Dec. 2009; also NASA/TM—2010-215676, Mar. 2010. 

31. Broeren, A.P., Whalen, E.A., Busch, G.T., and Bragg, M.B., “Aerodynamic Simulation of Runback 

Ice Accretion,” Journal of Aircraft, Vol. 47, No. 3, May-Jun. 2010, pp. 924-939. 

32. Busch, G.T., Broeren, A.P., and Bragg, M.B., “Aerodynamic Simulation of a Horn-Ice Accretion on a 

Subscale Model,” AIAA Paper 2007-0087, 45th Aerospace Sciences Meeting and Exhibit, Reno, NV, 

Jan. 2007. 

33. Busch, G.T., Broeren, A.P., and Bragg, M.B., “Aerodynamic Simulation of a Horn-Ice Accretion on a 

Subscale Model,” Journal of Aircraft, Vol. 45, No. 2, Mar.-Apr. 2008, pp. 604-613.