Andy P. Broeren

Glenn Research Center, Cleveland, Ohio

Jeff M. Diebold and Michael B. Bragg

University of Illinois, Urbana, Illinois

Aerodynamic Classifi cation of Swept-Wing

Ice Accretion

NASA/TM—2013-216381

May 2013

DOT/FAA/TC-13/21
NASA STI Program . . . in Profi le

Since its founding, NASA has been dedicated to the 

advancement of aeronautics and space science. The 

NASA Scientifi c 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 Offi cer. It collects, 

organizes, provides for archiving, and disseminates 

NASA’s STI. The NASA STI program provides access 

to the NASA Aeronautics and Space Database and 

its public interface, the NASA Technical Reports 

Server, 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 signifi cant phase 

of research that present the results of NASA 

programs and include extensive data or theoretical 

analysis. Includes compilations of signifi cant 

scientifi c and technical data and information 

deemed to be of continuing reference value. 

NASA counterpart of peer-reviewed formal 

professional papers but has less stringent 

limitations on manuscript length and extent of 

graphic presentations.

• TECHNICAL MEMORANDUM. Scientifi c 

and technical fi ndings 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. Scientifi c and 

technical fi ndings by NASA-sponsored 

contractors and grantees.

• CONFERENCE PUBLICATION. Collected 

papers from scientifi c and technical 

conferences, symposia, seminars, or other 

meetings sponsored or cosponsored by NASA.

• SPECIAL PUBLICATION. Scientifi c, 

technical, or historical information from 

NASA programs, projects, and missions, often 

concerned with subjects having substantial 

public interest.

• TECHNICAL TRANSLATION. Englishlanguage translations of foreign scientifi c and 

technical material pertinent to NASA’s mission.

Specialized services also include creating custom 

thesauri, building customized databases, organizing 

and publishing research results.

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 443–757–5803

• Phone the NASA STI Information Desk at

 443–757–5802

• Write to:

 STI Information Desk

 NASA Center for AeroSpace Information

 7115 Standard Drive

 Hanover, MD 21076–1320
Andy P. Broeren

Glenn Research Center, Cleveland, Ohio

Jeff M. Diebold and Michael B. Bragg

University of Illinois, Urbana, Illinois

Aerodynamic Classifi cation of Swept-Wing

Ice Accretion

NASA/TM—2013-216381

May 2013

DOT/FAA/TC-13/21

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135
Available from

NASA Center for Aerospace Information

7115 Standard Drive

Hanover, MD 21076–1320

National Technical Information Service

5301 Shawnee Road

Alexandria, VA 22312

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

Trade names and trademarks are used in this report for identifi cation 

only. Their usage does not constitute an offi cial endorsement, 

either expressed or implied, by the National Aeronautics and 

Space Administration.

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

Acknowledgments

The authors would like to thank Joe Botalla at the University of Illinois who wrote the fi rst draft of the literature review of 

swept wings with ice accretion and thus help initiate this paper. Abdi Khodadoust provided hard to acquire references from his 

extensive collection of icing papers. Several colleagues shared their expertise by reading a draft of this document and providing 

helpful comments and suggestions. This includes: Frederic Moens at ONERA; Cris Bosetti, Don Cook, Abdi Khodadoust, Adam 

Malone and Ben Paul at Boeing; Jim Riley at FAA; Sam Lee, Mark Potapczuk and Mary Wadel at NASA Glenn. The authors 

from the University of Illinois were supported in this effort by a grant from the FAA Technical Center, DOT_FAA 10-G-0004, 

with technical monitor Jim Riley. NASA’s Atmospheric Environment Safety Technologies Project of the Aviation Safety Program 

also supported this work.
NASA/TM—2013-216381 1 

Aerodynamic Classification of Swept-Wing Ice Accretion 

Andy P. Broeren 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Jeff M. Diebold and Michael B. Bragg 

University of Illinois 

Urbana, Illinois 61801 

Abstract 

The continued design, certification and safe operation of swept-wing airplanes in icing conditions rely 

on the advancement of computational and experimental simulation methods for higher fidelity results 

over an increasing range of aircraft configurations and performance, and icing conditions. The current 

state-of-the-art in icing aerodynamics is mainly built upon a comprehensive understanding of twodimensional geometries that does not currently exist for fundamentally three-dimensional geometries such 

as swept wings. The purpose of this report is to describe what is known of iced-swept-wing aerodynamics 

and to identify the type of research that is required to improve the current understanding. Following the 

method used in a previous review of iced-airfoil aerodynamics, this report proposes a classification of 

swept-wing ice accretion into four groups based upon unique flowfield attributes. These four groups are: 

ice roughness, horn ice, streamwise ice, and spanwise-ridge ice. For all of the proposed ice-shape 

classifications, relatively little is known about the three-dimensional flowfield and even less about the 

effect of Reynolds number and Mach number on these flowfields. The classifications and supporting data 

presented in this report can serve as a starting point as new research explores swept-wing aerodynamics 

with ice shapes. As further results are available, it is expected that these classifications will need to be 

updated and revised. 

Nomenclature 

AR aspect ratio 

c chord length 

Cl sectional lift coefficient 

Cd sectional drag coefficient 

CD wing drag coefficient 

CL wing lift coefficient 

Cl,max maximum sectional lift coefficient 

CL,max maximum wing lift coefficient 

Cm sectional pitching-moment coefficient 

CM wing pitching-moment coefficient 

cmac mean aerodynamic chord 

Cp pressure coefficient

k roughness height 

M Mach Number 

rle leading-edge radius 

Re Reynolds Number 
NASA/TM—2013-216381 2 

s surface distance measured from leading edge 

V velocity 

x chordwise coordinate 

y spanwise coordinate 

z vertical coordinate 

α angle of attack 

 Taper ratio 

Λ sweep angle 

FAA Federal Aviation Administration 

IRT Icing Research Tunnel 

LWC Cloud liquid water content 

MVD Median volumetric diameter 

NACA National Advisory Committee for Aeronautics 

NASA National Aeronautics and Space Administration 

ONERA Office National d Etudes et de Recherches Aérospatiales 

SLD Supercooled large droplet 

1.0 Introduction 

Ice accretion and its aerodynamic effect on highly three-dimensional swept wings are extremely 

complex and important to the continued design, certification and safe operation of small and large 

transport aircraft. There is increasing demand to balance trade-offs in aircraft efficiency, cost and noise 

that tend to compete directly with allowable performance degradations over an increasing range of icing 

conditions. Aircraft icing research has now reached the level of maturity that computational methods and 

experimental tools are currently being used to address many of these challenges. However, knowledge 

gaps do remain for swept-wing geometries and larger droplet icing conditions. In prior research on iced 

airfoils it was extremely helpful to frame the problem by identifying classes of ice shapes based on their 

effect on the flowfield. The purpose of this paper is to propose similar ice-shape classifications for the 

three-dimensional swept-wing case. As this paper will present, this is extremely challenging at this point 

in our knowledge of these flowfields. Thus, this initial proposed classification is intended as a guide for 

future research and one that is anticipated to evolve as we learn more about swept-wing ice accretion and 

their flowfields. 

1.1 Ice-Shape Classification for Airfoils 

Systematic and parametric studies of two-dimensional airfoil icing aerodynamics have provided much 

information about aerodynamic performance and associated flowfields. Research documenting iceaccretion geometry and aerodynamic performance has been conducted since the 1940s. Since the 1980s 

researchers have also focused on numerical methods for ice-accretion prediction and the resulting 

aerodynamic performance. Also in this time period experimentalists have provided more flowfield details 

and thus a better understanding of the physics of the ice-accretion flowfield on airfoils and straight wings. 

This research is summarized in the review articles by Lynch and Khodadoust (Ref. 1) where the icing 

aerodynamic performance is examined and by Bragg, Broeren and Blumenthal (Ref. 2) where the 

flowfield characteristics are reviewed and ice-shape classifications are presented. Effects from airfoil 

geometry, ice-accretion geometry, and size and location of the ice shape are currently understood along 

with the fundamentals of the flowfield. Thus, the underlying physics that drive the performance 

degradation is also known. Simulation of ice accretion and the Reynolds number effects are understood. 
NASA/TM—2013-216381 3 

These accomplishments have led to, and continue to lead to, better design, certification, and modeling of 

aircraft with ice accretion and thus improved safety. These accomplishments in two-dimensional airfoil 

icing aerodynamics were made primarily through systematic and fundamental experimental studies. 

There are many types of ice accretion classified by the physics of the ice accretion process and these 

labels are more familiar to many engineers. These include rime ice, glaze or clear, ice, mixed ice, beak 

ice, runback ice, intercycle ice and many others. Although this nomenclature is appropriate for the 

accretion of ice, they may not be as useful when the objective is aerodynamic simulation. Bragg et al. 

(Refs. 2 and 3) examined the icing aerodynamics literature and developed four fundamental types based 

on the flowfield physics. These ice-shape classifications are (1) roughness, (2) horn ice, (3) streamwise 

ice, and (4) spanwise-ridge ice. The important aerodynamics associated with each of these is summarized 

in the following paragraphs. Much of this discussion is taken from Bragg et al. (Ref. 2) and more details 

can be found in the original references. 

1.1.1 Ice Roughness 

Ice roughness usually occurs during the initial stages of the accretion process before a significant ice 

shape, such as a horn, is accreted. The other three ice types are also “rough,” but here the accreted 

roughness does not significantly alter the airfoil contour nor the inviscid flowfield. A key aerodynamic 

feature of ice roughness is that it is usually much larger than the local boundary-layer thickness even at 

the very early stages of development. These roughness elements act as bluff bodies with their own threedimensional separated flowfield. In fact, the distinguishing characteristic of ice roughness and the other 

types of ice can be described in terms of the extent of boundary-layer separation. The separated flow 

regions generated by ice roughness are fundamentally three-dimensional, are very local to the roughness 

elements, and are similar in scale to the roughness itself. 

Roughness is characterized by its height, concentration (or density), surface location and distribution. 

Its effect on airfoil performance depends on all of these parameters. Roughness element shape can also be 

significant, but the irregular shapes seen in ice roughness are not thought to be as important as, and are 

certainly less well understood, than the other parameters. Roughness can affect airfoil performance in 

several ways. The roughness elements themselves can extract momentum from the flow and cause 

boundary-layer separation near the trailing edge (Refs. 3 and 4). Trailing-edge separation can contribute 

to reduced lift coefficients and reduced stalling angle of attack. Roughness usually promotes bypass 

transition in the boundary layer that can lead to increased skin friction. These effects manifest themselves 

globally as performance degradation—increased drag and decreased maximum lift. 

1.1.2 Horn Ice 

Longer exposures to icing conditions conducive to the formation of glaze and mixed ice usually result 

in the formation of horn ice. The horn-ice shape can be characterized by its height, the angle it makes 

with the chord line, and its location indicated by s/c. The most distinctive feature of the horn-ice flowfield 

is the large separation bubble that forms downstream of the horn. The stagnation point is usually located 

on the ice shape, and the boundary layer cannot negotiate the large adverse pressure gradient encountered 

at the tip of the horn. In some cases, such as beak ice, the stagnation point may still be located on the 

airfoil upstream of the horn. In any case, for horn ice, the separation location remains approximately fixed 

at the horn tip over a large angle-of-attack range. The separated shear layer undergoes transition to 

turbulent flow and then usually reattaches to the airfoil surface downstream. The separation bubble causes 

a large redistribution of pressure that results in pitching moment changes and decreased lift; it also greatly 

increases the airfoil drag. This flowfield is also known to be unsteady and three-dimensional. These 

factors make computational simulation challenging. Therefore, understanding the behavior of the 

separation bubble is key to understanding horn-ice aerodynamics. 

For this shape, it is somewhat paradoxical that the separation bubble features that make the flow 

challenging for computational modeling simplifies the geometric simulation required to reasonably 

reproduce the effects on lift, drag, and pitching moment. For example, several studies (Refs. 5 to 9) have 
NASA/TM—2013-216381 4 

shown that very simple geometries, such as a leading-edge “spoiler,” can reproduce the performance 

characteristics of a horn-ice shape over a large angle of attack range. A simple geometry representing the 

height, angle, and location of the ice horn, essentially generates an equivalent separation bubble on the 

airfoil and, hence, very similar performance results. Also, the addition of surface roughness is shown in 

some studies (Refs. 6, 8 to 11) to have only minor effects on the integrated performance. 

1.1.3 Streamwise Ice 

Streamwise ice has the smallest effect of the non-roughness ice accretions and has received the least 

attention in the literature. Therefore, an understanding of the aerodynamics of these accretions is not as 

developed as for the horn and spanwise-ridge ice. Streamwise ice is often formed as a result of rime icing 

conditions that occur at cold temperatures when the incoming water droplets freeze on the surface upon 

impingement. As a result, the ice accretion follows the contour of the airfoil surface. For the streamwise 

ice geometries that are conformal to the airfoil leading edge, the stagnation point at moderate lift 

coefficients occurs on the ice shape and the boundary layer remains attached as it flows around the 

leading edge of the ice and downstream on the upper surface. Since the streamwise ice/airfoil intersection 

is not smooth, an adverse pressure gradient may exist in this area and flow separation may occur in the 

junction region. The flow separation location is not fixed to a specific point on the ice shape as it is fixed 

to the tip for a horn shape, but can move upstream or downstream depending upon the angle of attack and 

the state of the incoming boundary layer, which depends on the surface roughness, Reynolds number, and 

other factors. 

At large exposure times, or when the icing conditions are otherwise appropriate, streamwise ice 

shapes can occur that are not as conformal to the original airfoil surface and may grow a horn-like feature 

into the flow. In some cases, flow separation may result as with horn ice. These separation bubbles are 

typically much smaller than for horn ice and thus do not have as large of an effect on the stall mechanisms 

(Refs. 10 and 12). This means that other flowfield features, such as trailing-edge separation, play a 

significant role in the aerodynamics. For streamwise ice, the specific ice geometry and surface roughness 

characteristics can be important factors in the aerodynamics. 

1.1.4 Spanwise-Ridge Ice 

Spanwise-ridge ice is perhaps most often associated with supercooled large-droplet (SLD) icing 

conditions. Usually these accretions form downstream of leading-edge thermal or mechanical iceprotection systems and, despite their association with SLD, can occur for all drop-size ranges. Runback 

icing can form ridge accretions which usually occur when there is a heated leading-edge ice-protection 

system that is not removing all of the impinging water. Some water flows downstream on the surface 

from the heated section and freezes on the cooler, unheated surface. Because of the formation mechanisms 

just described, ridges often exhibit extensive spanwise variation in their geometry. These properties, and the 

associated flowfield, make the spanwise-ridge-type accretion different from the horn shapes discussed in 

Section 1.1.2. Spanwise-ridge ice can have more severe aerodynamic effects, and the shapes themselves are 

typically more three-dimensional than horn shapes in terms of gross geometry or spanwise variation. 

Spanwise ridges are generally located farther downstream than horns or streamwise shapes. This distance 

allows the boundary layer to develop, perhaps transition, or become transitional because of small ice 

roughness upstream of the spanwise-ridge. Thus the spanwise ridge acts as a flow obstacle. 

The unique characteristics of the ridge—large spanwise variation and downstream location—make 

the resulting aerodynamics and performance different from those for horn ice. The chief similarity is the 

large separation bubble downstream of the spanwise ridge, which can be complex, three-dimensional, and 

highly unsteady. Spanwise ridges also have a smaller separation bubble upstream of the ridge because the 

ridge is located well downstream of the stagnation point. This upstream separation can be a further 

challenge for computational simulation. The fact that the ridge is located downstream of the stagnation 

point also amplifies the effect of airfoil geometry, which is characterized by the clean airfoil pressure 
NASA/TM—2013-216381 5 

distribution; since spanwise ridges are located well downstream of the stagnation point. Other important 

features are the geometry, size, and location of the ridge on the airfoil (Ref. 13). 

Later research conducted for spanwise ridges has revealed some potentially important differences in 

the fundamental aerodynamics between “short” and “tall” ridges. Broeren et al. (Refs. 14 to 16) proposed 

a subclassification that distinguishes short and tall spanwise-ridge shapes on the basis of their effect on 

the airfoil pressure distribution and resulting aerodynamics. The tall-spanwise-ridge characteristics were 

described in the previous paragraph. In the case of short ridges, the separation bubble is small and stable 

and does not substantially increase in size with increasing angle of attack. The effect on the pressure 

distribution is minimal, thus resulting in significantly lower aerodynamic penalties relative to tall 

spanwise ridges. Although much has been learned from this recent work, more research is required to 

further develop the short spanwise ridge aerodynamic characteristics. 

1.1.5 Summary 

Sections 1.1.1 to 1.1.4 briefly describe the aerodynamic classification of ice accretion on airfoils. Real 

ice accretion may not fit neatly into only one classification, and of course, it is not realistic to think of 

these classifications as rigid, absolute or unchanging. Undoubtedly more research will yield greater 

insight into iced-airfoil aerodynamics and may lead to further development of these concepts, as has 

already occurred in the case of spanwise-ridge ice. Future considerations notwithstanding, the present 

classification system developed for iced airfoils formed the foundation for the research described in this 

report extending these concepts to three-dimensional geometries. 

1.2 Objectives and Challenges 

NASA, FAA, ONERA, the University of Illinois, Boeing and others are embarking on a research 

program whose goals are to improve our understanding and ability to model ice accretion and the 

resulting aerodynamic effect on full-scale, three-dimensional swept wings. This presents numerous 

challenges for both computational and experimental research as we extend our two-dimensional models 

and techniques where possible and develop new approaches and techniques to address the more complex 

three-dimensional nature of the problem. 

The parameter space for the three-dimensional wing case is much larger and more complex than for 

airfoils. In addition to airfoil geometry, ice-shape geometry, size and location, the three-dimensional wing 

geometry must be considered. Here the airfoil and ice accretion vary spanwise and wing sweep angle, 

twist, taper, dihedral and aspect ratio must be considered. Given the vast range of swept-wing 

configurations, geometries and airfoil sections, it is useful to define and limit the scope of consideration 

for this fundamental study. The application of this research is directed toward commercial jet transport 

airplanes including: business, regional, single-aisle and wide-body types. Therefore, the range of wing 

sweep considered is 25 to 40. In terms of ice accretion physics and aerodynamics, sweep becomes an 

increasingly important parameter for values greater than 30. The appropriate range of aspect ratio is from 

about 6 to 9. Although leading- and trailing-edge devices are standard equipment on airplane wings, the 

fundamental studies presented in this review tend to have hard leading edges and fixed trailing edges. The 

effects of wing-mounted engines, pylons and side-of-body fuselages are also not considered here. 

When the ice-shape classifications for airfoils were developed there was a larger and better 

understanding of the performance and aerodynamics than currently exists for three-dimensional iced 

wings. While several high-quality experimental programs have been conducted on three-dimensional 

wings with ice, these studies were focused on specific ice accretions or geometries and provide only a 

piece of the puzzle. In addition, most of these experiments were conducted at low-Reynolds number and 

the Reynolds number effects for iced three-dimensional wings are not yet understood. Similar 

experiments were initially conducted on iced airfoils and were useful, but did not provide the data needed 

for a more in depth understanding of the aerodynamics. Real progress toward the ice-shape classifications 

was made when systematic experimental studies that included ice-shape geometry, airfoil geometry, 

Reynolds and Mach number effects were conducted. A corresponding systematic study, analogous to the 
NASA/TM—2013-216381 6 

two-dimensional research, is needed to provide major advances in our understanding of three-dimensional 

iced-wing aerodynamics. Therefore, this research will be conducted within the framework of threedimensional ice-shape classification to increase its efficiency and productivity. The challenge is, of 

course, that the knowledge required to formulate the classifications is very limited. As research is 

conducted and the knowledgebase increases, the classifications will undoubtedly change and improve. 

Thus, with the understanding that this will be the first of probably many versions of a threedimensional, swept-wing, ice-shape classification system, the objective of this paper is to present this 

initial version. The methods used in developing aerodynamic classifications of ice accretion on airfoils 

are applied to ice accretion on swept wings. Given what information is known, the approach is: 

 To review the fundamental stalling mechanisms and associated flowfield characteristics of swept 

wings. 

 To review the physical characteristics of ice accretion observed on swept wings. 

 To review aerodynamic characteristics of iced-swept wings in the framework of two-dimensional 

airfoil classifications. 

 To describe the type of additional data needed to better define the classification system for swept 

wings. 

Undoubtedly the classification system will require many improvements and changes, but it is hoped that 

this document provides the initial framework for improved understanding of the important and complex 

problems and is ultimately an important tool for both researchers and practicing engineers involved in 

aircraft icing. 

2.0 Background: Stalling Characteristics and Ice Accretion on Swept Wings 

Before embarking on a detailed review of iced-swept-wing aerodynamics, it is appropriate to identify 

the key features of the flowfield on the clean swept wing. Emphasis is placed on the stalling 

characteristics since this tends to be of primary importance in aircraft safety-related icing studies. It is 

well known that a major effect of ice accretion is to impact the stalling characteristics of a wing, thereby 

limiting the maximum lift coefficient to values that are often well below that of the clean wing. This 

review and the initial ice-shape classification system focus on stalling characteristics because this tends to 

be the most salient feature delineating the aerodynamics of the various types of ice shapes. This 

background discussion also reviews the physical characteristics of ice accretion on swept wings. These 

characteristics can be significantly different from straight-wing ice accretion due to the effects of 

spanwise flow on swept wings. It is also useful to identify these physical characteristics is terms of the 

icing conditions which are more commonly used to describe ice accretion. Then, in Section 3.0, the 

swept-wing ice accretion is described in terms of the associated aerodynamics as has been accomplished 

for two-dimensional airfoils. 

2.1 Aerodynamic Characteristics of Swept-Wing Stall at Low-Mach Number 

By studying the stalling characteristics of swept wings in the baseline, uniced condition, parallels may 

be drawn to the stalling characteristics with accreted ice. The technical literature describes effects due to 

leading-edge radius, Reynolds number and surface roughness. All of these factors can usually be related 

to some characteristics or attributes of leading-edge ice accretion. Therefore, this review yields insight 

into the general flowfield characteristics of importance for swept wings with hard leading edges. 

Furlong and McHugh (Ref. 17) reviewed the stalling characteristics of swept wings and identified 

leading-edge and trailing-edge separations as key, but distinct, contributors to the wing stall. As the 

names imply, these basic characteristics are fundamentally different, and either type or a combination of 

the types may occur on a swept wing. When trailing-edge separation dominates, the spanwise outflow 
NASA/TM—2013-216381 7 

leads to an excessively thick boundary layer on the outboard portions that is more susceptible to 

separation. At the same time, the inboard sections become more resistant to trailing-edge separations. 

(This is true for a plain swept wing, e.g., no side-of-body effects.) When leading-edge separation 

dominates, a leading-edge separation bubble and spanwise pressure gradient result in a spanwise-running, 

spiral, leading-edge vortex. This vortex runs outward along the leading edge until it leaves the wing 

surface at a location that moves inboard along the trailing edge with increasing angle of attack. It is 

possible to have a combination of these features and the type of separation is a function of several 

variables including leading-edge radius, sweep angle, Reynolds number and aspect ratio. The formation 

of the leading-edge separation bubble on the individual airfoil sections is sufficient, but not necessary for 

the formation of the leading-edge bubble on the swept wing. Other effects such as induced camber, or ice 

accretion, may cause a leading-edge separation bubble to form. 

Poll (Ref. 18) performed a fundamental flow visualization study on swept wings to investigate the 

formation and development of the spanwise-vortex resulting from leading-edge separation. Variables 

considered were sweep angle, leading-edge radius and Reynolds number. For the 30, swept, rectangular 

wing having a sharp leading edge (rle/c = 0.0003) at Re = 1.7×106

, Poll noted the formation of a full-span, 

leading-edge, spiral vortex beginning at 3.5 angle of attack. The leading-edge vortex persisted up to α = 

9, above which it burst leaving a leading-edge vortex on the inboard portion and separated flow over the 

outboard portion. Further increases in angle of attack simply caused the spanwise extent of the vortex to 

decrease toward the root. The description of the stall characteristics matches the leading-edge separation 

type identified by Furlong and McHugh. Variation in Reynolds number from 0.75×106

 to 3.0×106 had no 

significant effect on the flowfield for the sharp leading-edge model. 

When the leading-edge radius was increased to rle/c = 0.012, a short leading-edge separation bubble 

was observed at α = 6. The full-span separation bubble persisted up to 12 angle of attack where 

evidence of reverse flow over the tip was observed. Further increases in angle of attack resulted in more 

flow separation over the outboard portion. Poll suggested that there was a leading-edge vortex above the 

surface on the inboard portion of the wing. As before, changes in Reynolds number between 0.87×106

 and 

1.7×106

 had no significant effect on the flowfield. 

When the leading-edge radius was further increased to rle/c = 0.03 at Re = 1.7×106

, there was no 

evidence of a leading-edge separation bubble up to α = 15. Instead, Poll observed an inboard, leadingedge vortex and separated flow over the outboard portion of the wing. In this case, it appears that the stall 

type is more similar to the trailing-edge type identified by Furlong and McHugh. Further increases in 

angle of attack resulted in a similar flowfield with the vortex moving closer to the root. For this leadingedge radius, lowering the Reynolds number to 0.9×106

 resulted in significant flowfield differences, with 

large regions of separated flow on the forward portion of the wing increasing in chordwise extent to the 

tip region. A small leading-edge vortex was observed very close to the wing root. 

In Poll’s experiments on the 30 swept wing, the evolution of the spanwise-running, leading-edge 

vortex was related to the leading-edge separation. These flowfield features were also observed in a lowReynolds number study of a 3.5 percent scale Generic Transport Model (GTM) (Ref. 19). In this case, the 

wing sweep was 25 and the Reynolds number was 0.24×106

. Surface-oil flow visualization was 

performed on the baseline configuration for several angles of attack documenting the time-averaged 

surface flowfield on the wings leading up to stall. The flow visualization image of the right wing in 

Figure 1 for α = 8 shows the formation of a spanwise-running, leading-edge vortex on the surface of the 

wing approximately parallel to the leading edge. This complex flowfield is also described with the aid of 

the schematic representation adapted from Poll. Poll refers to the leading-edge separation location as the 

“primary” separation (Fig. 1) that led to a vortex sheet rolling up to form a strong vortex very close to the 

wing surface. This vortex drew free-stream fluid down to the surface along the reattachment line, thus 

forming a leading-edge separation bubble with surface flow moving upstream between the reattachment 

location and the leading edge. 

 
NASA/TM—2013-216381 8 

Poll’s model suggests that initially this reverse flow experiences a favorable pressure gradient, but 

then moves into a region of adverse pressure gradient leading to the “secondary” separation of the reverse 

flow. This is indicated in Figure 1 and can be seen in the line of oil accumulation upstream of and 

approximately parallel to the reattachment line. Downstream of the reattachment line, the surface flow 

was uniform and generally in the streamwise direction to the wing trailing edge. While this flowfield 

description is analogous to the nominally two-dimensional leading-edge separation bubble, the coherent 

spiral vortex is clearly a significant three-dimensional structure. The leading-edge vortex is the key 

flowfield characteristic in this case. While it may be technically accurate to continue to refer the flowfield 

in Figure 1 in leading-edge separation bubble terminology, such terminology lacks the key contribution of 

the spanwise flow. Therefore, in this report this flowfield is described in terms of a leading-edge vortex 

that lies close to the wing surface. 

This description of leading-edge separation leading to the swept-wing stall is consistent with the 

description of leading-edge separation described by Furlong and McHugh (Ref. 17), who refer to the 

leading-edge vortex as a key flowfield characteristic. Further increases in angle of attack cause the 

leading-edge vortex to leave the surface of the wing and turn downstream resulting in separated flow over 

the outboard portion of the wing. This scenario is depicted in Figure 2 for the 3.5 percent GTM wing with 

the angle of attack increased to 12. The presence of the wing-mounted engine and pylon caused both an 

inboard and outboard leading-edge vortex to form. Both vortices appear to leave the surface, but on the 

inboard portion, the surface flow remains attached to the trailing edge. On the outboard portion, there was 

a large region of reverse flow on the surface indicative of boundary-layer separation. 

Research has shown that sweep angle and leading-edge radius, or more generally, leading-edge 

geometry, both affect the influence of Reynolds number on aerodynamic performance. Polhamus 

(Ref. 20) provides a simple example in Figure 3 that shows the declining effect of Reynolds number on 

maximum lift with increasing sweep angle. As the sweep angle increases above 40, the influence of the 
NASA/TM—2013-216381 9 

vortex flow increases thus mitigating Reynolds number effects. For more modest sweep angles around 

30, both Reynolds number and vortex flow effects are important in the stall characteristics for the wing 

shown in Figure 3. Polhamus suggests that the leading-edge vortex flow (presumably similar to that 

observed by Poll and summarized by Furlong and McHugh) may play a role in the stall behavior for the 

30 sweep configuration in Figure 3. 
NASA/TM—2013-216381 10 

In addition to sweep-angle effects, Poll showed that for sharp or small radius leading edges, the 

flowfield was independent of Reynolds number over the range tested. Furlong and McHugh also reached 

a similar conclusion based upon data at Reynolds number up to 6.0×106

. This is shown in Figure 4 for 

inflection lift coefficient (or “usable” lift coefficient—designating the lift coefficient at which there is a 

shift in the wing aerodynamic center (Ref. 17)) variation with Reynolds number on the wing composed of 

circular-arc airfoil sections that had a sharp leading edge. Similar results were observed on the wing 

composed of NACA 641-112 airfoil sections with roughness. The application of roughness on the leading 

edge induced a leading-edge vortex flow at lower lift coefficient than in the smooth leading-edge 

configuration, thus mitigating the beneficial effects of increasing Reynolds number. Neely and Connor 

(Ref. 21) also observed similar effects of leading-edge roughness as shown in Figure 5. These results are 

not unexpected. Fixing the point of separation (say with ice accretion) has limited the effects of Reynolds 

number in myriads of airfoil tests. Therefore, any leading-edge configuration, such as a sharp leading 

edge, roughness, or perhaps ice accretion, may work in concert with the leading-edge vortex to mitigate 

Reynolds number effects on the stall behavior. 
NASA/TM—2013-216381 11 

This brief review has identified key flowfield features for swept wings such as leading-edge 

separation and leading-edge vortex formation. These features govern the wing stalling process and may 

be influenced by leading-edge geometry (including high-lift systems), sweep angle, aspect ratio, Reynolds 

number and other parameters. In many cases, ice accretion can be thought of as a modification to the 

leading-edge geometry. An important exception is the case of spanwise-ridge ice where the leading edge 

may be free of large ice accretion. For leading-edge ice accretion on airfoils, the boundary-layer 

separation point is often fixed on the ice shape independent of angle of attack or Reynolds number. It is 

reasonable to assume that in some cases, ice accretion will also determine the separation location for 

swept wings. Questions then arise regarding the formation and evolution of the leading-edge vortex. 

Previous studies indicate that this formation and evolution will govern the iced-wing aerodynamics. So 

understanding this important characteristic is a key to understanding the iced-wing flowfield. 

The works cited here were fundamental studies that employed basic plain wings without high-lift 

systems and other components of real wings. Lynch and Khodadoust (Ref. 1) point out that the proper 

assessment of three-dimensional icing effects relies on a sound understanding of the spanwise variation in 

stall initiation on the baseline, uniced wing. They note that well designed wings do not normally have the 

stall occur simultaneously across the span as for two-dimensional airfoils. Typically, there is a particular 

spanwise location that is critical to stall initiation. The addition of wing-mounted engines, nacelles and 

pylons can also modify the stall behavior. Therefore, it is important to know the critical section, span-load 

distribution, baseline wing stalling characteristics and how ice accretion affects the stall progression 

relative to the critical section. 

2.2 Characteristics of Swept-Wing Ice Accretion 

Vargas (Ref. 22) provides an excellent historical and technical summary of swept-wing ice accretion. 

He notes that for rime ice accretion, there is no fundamental difference in the accretion characteristics for 

any sweep angle in the range of interest here. Consider the example shown in Figure 6 on a GLC-305 

wing having 28 leading-edge sweep taken from Vargas et al. (Ref. 23) Typical of rime ice accretion, 

there is a smooth zone along the attachment line and large roughness and rime feather formations 

downstream on the lower surface. The cross-sectional pictures show that the rime shape in this case 

consists of a protuberance oriented into the free-stream flow direction. This could be thought of as a 

leading-edge extension. Another rime ice example is shown in Figure 7 on a NACA 0012 wing with 45

sweep angle. Note that the “smooth zone” is not entirely smooth in the region of the attachment line. In 

this case, the strong spanwise flow associated with the larger sweep angle led to this non-smooth region. 

This ice accretion is also characterized by large rime feathers that are also observed on airfoils subjected 

to similar conditions. Based upon observations of airfoil ice accretion it is also possible that the rime 

shape may be more conformal to the leading edge for other cloud conditions. 

For glaze icing, increasing wing sweep has a profound effect on the characteristics of the resulting 

accretion. Depending upon the sweep angle, air temperature, cloud liquid water content and other factors, 

highly three-dimensional features called “scallops” or “lobster tails” may form. Through several research 

efforts, Vargas (Ref. 22) has identified sweep angle and icing conditions leading to: no scallops, 

incomplete scallops and complete scallop glaze and mixed ice accretion. Illustrations of each shape can be 

seen in Figure 8. An example of a complete scallop ice shape is shown in Figure 9. Note the regular gaps 

in the glaze “horns” in the spanwise direction. Glaze ice accretion with typical “double horn” features can 

also incorporate the scallop formations shown in Figure 9. 

In his review, Vargas (Ref. 22) summarizes many important observations regarding the formation of 

scallop tips. Scallop formations are composed of glaze-ice feathers that grow from roughness elements 

deposited in the initial stages of ice accretion. Ice feathers are individual ice features that are attached to a 

wing surface or main ice shape at discrete points and fan outward as they grow. Feathers that form 

scallops grow much larger in a “preferred” direction that is perpendicular to the local free-stream flow 

direction. Thus they tend to “fan out” in a spanwise direction while being relatively thin in the streamwise 

direction. Vargas has developed the definition of “critical distance” with regard to scallop formation. 
NASA/TM—2013-216381 12 

 
NASA/TM—2013-216381 13 

 
NASA/TM—2013-216381 14 

The “critical distance” is the surface length measured from the attachment line to the boundary where the 

roughness elements become feathers during the initial stages of the accretion process. When the critical 

distance approaches zero, the feather formation dominates and the feathers grow into scallop formations. 

The main categories of rime and glaze ice accretion can also be applied to the initial ice roughness 

that forms during the first 1 to 2 min. of icing exposure. The characteristics can differ significantly 

depending upon whether the roughness is rime, glaze or mixed. Figure 10 compares glaze and rime 

roughness formation on a swept-wing model. The fundamental features of the roughness are essentially 

identical to initial ice roughness observed on straight wings and airfoils. There are some differences in the 

region of the attachment line. Particularly noticeable for the glaze roughness is the lack of a smooth zone 

typically observed on straight wings and airfoils. In the case of Figure 10, the roughness is observed from 

the attachment line downstream to the ice accretion limit. Although not readily seen in the rime roughness 

image, the region of the attachment line was not entirely smooth. 

Another type of ice accretion is due to the operation of ice-protection systems. Deicing systems that 

employ cyclic operation can leave behind residual ice and form intercycle ice between the cycles of the 

deicer. Usually, the resulting residual and intercycle ice accretion are not fundamentally different from 

those found on straight wings or airfoils. Anti-icing systems provide continuous heating to the leadingedge impingement region. Any incoming water that does not evaporate flows downstream and freezes 

where the surface temperature drops below 0 °C. In this condition a “runback” ridge or ice formation may 

form. Some examples of these types of ice accretions can be found in References 24 and 25. 

These many variations in type, size, location and roughness level of ice accretion support the notion 

of aerodynamic classification for improved simulation. That is, if a small number of ice-shape groups can 

be identified having similar fundamental aerodynamic characteristics, then the exact nature or origin of 

the ice accretion itself becomes less critical in terms of aerodynamics. This approach was successful for 

ice accretion on airfoils. The remainder of this report describes a similar approach for swept wings. 
NASA/TM—2013-216381 15 

3.0 Aerodynamic Classification of Swept-Wing Ice Accretion 

The purpose of this section is to review the existing body of technical literature on iced-swept wing 

aerodynamics in the context of a classification method based upon fundamental aerodynamic 

characteristics. The classification method used in this report closely follows the previous two-dimensional 

ice-shape classification (Ref. 2). Thus classification is based on the fundamental flow physics and ice 

geometries that generate these effects on swept-wing aerodynamic performance. However, the ice-shape 

classification is limited to geometries that are known to form in icing conditions as previously discussed 

in Section 2.2. The four main classes of swept wing ice accretion are: ice roughness, horn ice, streamwise 

ice and spanwise-ridge ice. Each classification is described in the sections below that focus on the unique 

aerodynamic characteristics as they are currently understood. Where possible, the aerodynamic 

performance data are interpreted in terms of the key flowfield features observed in detailed studies. 

Where information is lacking, there is a description of the research needed to provide for a better 

understanding of the iced-wing aerodynamics. 

An additional limitation of the following review and proposed classification scheme is the lack of a 

detailed discussion of the spanwise variation of the ice accretion. Modern highly twisted and tapered 

swept wings have very different flow and icing conditions from the root to the tip. Under certain 

conditions this can lead to large variation in ice shape geometries along the span. The effect of this 

variation also depends on span-load distribution of the wing. Lynch and Khodadoust (Ref. 1) note in their 

review that it is important to account for the spanwise variation in stall initiation with a certain station 

being most critical. Thus ice accreted at this critical location can cause an immediate effect on the wing 

stall, while ice at a different spanwise station may have to be more significant before it would become the 

critical stall initiation location. The message here is that all leading edge ice is not equal on a swept wing 

and that some spanwise locations are more sensitive in terms of the degradation in wing aerodynamic 

performance than others. Thus our approach in this paper is to extend the two-dimensional classifications, 

well supported by experimental data, to the three-dimensional swept-wing case guided by the limited 

three-dimensional data available. Still, as noted by Lynch and Khodadoust in 2001, the existing threedimensional data have not explored the effects of spanwise variation in ice accretion and the influence on 

the critical spanwise station. This presents an additional limitation in the resulting classification, and 

another area where more research is needed and future improvement is expected. 

3.1 Ice Roughness 

The roughness classification is based upon roughness associated with the initial stages of in-flight ice 

accretion. As described in Section 2.2, the characteristics of initial ice roughness on swept wings are 

fundamentally the same as on straight wings and airfoils (also Fig. 10). The possible exception is that the 

region of the attachment line may not always be smooth for swept wings. Ice roughness can be 

characterized by its height, location, chordwise extent, concentration and shape (Ref. 2). These effects are 

understood for airfoils, but much less so for swept wings. With regard to height, Vargas (Ref. 22) noted 

that “The roughness elements observed and measured...were much larger than the typical roughness 

elements considered in boundary-layer stability studies, and it is expected that they are of the order of the 

boundary-layer thickness or larger.” Such was also the case for ice roughness on airfoils that was much 

larger than the local boundary-layer thickness. Vargas goes on to state that the effects of roughness of this 

size on boundary-layer instabilities and transition are unknown. It is reasonable to assume that the local 

separated flowfield in the vicinity of the roughness elements is not fundamentally different than in the 

airfoil case. For this, Bragg et al. (Ref. 2) provide a detailed review of the roughness flowfield including 

the relevant phenomenological features, effects on transition and turbulence and impact on aerodynamic 

performance. For swept wings, the questions that arise have to do with the interaction of this flowfield 

with the spanwise and vortex flows. 

 
NASA/TM—2013-216381 16 

The National Advisory Committee for Aeronautics (NACA) conducted numerous experimental 

investigations into the aerodynamics of swept-wing configurations. Many of the studies explored 

Reynolds and Mach number effects and included cases with carborundum grains applied to leading edges. 

These studies provide a basis for aerodynamic effects, but do not contain detailed flowfield information. 

Koven and Graham (Ref. 26) conducted aerodynamic experiments in the NACA Langley 19-ft pressure 

tunnel with a model having an aspect ratio of 6, 37 leading-edge sweep, mean aerodynamic chord of 

43.55 in. and a taper ratio of 0.5 over a Reynolds number range of 2.0×106

 to 9.35×106 with 

corresponding Mach number range of 0.08 to 0.18. The main objective of this work was to investigate 

stall-control devices on the leading and trailing edges. Koven and Graham (Ref. 26) showed that a halfspan leading-edge slat positioned on the outboard portion of the wing effectively eliminated tip stall on 

the untwisted wing. Their flow visualization results were corroborated by the wing pitching-moment data 

that broke in a nose-down direction at stall. The authors also investigated the addition of full-span, 

double-slotted flaps in addition to the leading-edge slat. The lift, drag and pitching moment characteristics 

are shown in Figure 11 with the effect of roughness at Re = 6.8×106

 and M = 0.13. The roughness 

consisted of k/cmac = 0.00025 diameter carborundum grains applied between x/c = 0.02 on the lower 

surface to x/c = 0.08 on the upper surface on both the slat and inboard sections of the main wing leading 

edge. The roughness had a significant impact in CL,max reduction with very little increase in drag. The 

latter most likely occurred because of the larger increment in drag (both profile and induced) due to the 

leading-edge and trailing-edge devices. Koven and Graham data for identical roughness applied to the 

plain wing showed a significant drag penalty. The pitching moment data show an unstable (i.e., nose-up) 

break in the pitching moment at stall. However, Koven and Graham attribute this to the pitching-moment 

behavior of the multi-element airfoil section and not to tip stall as occurred for the plain wing, or the wing 

without the leading-edge slat. Therefore, in this case, it appears that the roughness did not affect the 

general stall pattern on the wing, but only the angle of attack and subsequent values of lift and drag 

coefficients. 

Neely and Conner (Ref. 21) also investigated leading-edge roughness effects in the Langley 19-ft 

pressure tunnel. Their model was a full-span wing with aspect ratio 4, 42 leading-edge sweep, mean 

aerodynamic chord = 34.71 in., and taper ratio = 0.625. Aerodynamic testing was conducted over a 

Reynolds number range of 1.7×106

 to 9.5×106 with corresponding Mach number range of 0.10 to 0.22. 
NASA/TM—2013-216381 17 

Flow visualization showed that the clean (untwisted) wing exhibited asymmetric tip stall. The stall 

progression began with separated flow beginning near the trailing edge of the outboard portions and 

progressing forward and inboard on the wing. In contrast, with k/cmac = 0.00032 diameter carborundum 

grains applied between x/c = 0.08 on the lower surface to x/c = 0.08 on the upper surface covering 

approximately 5 to 10 percent of the area, flow separation began at the leading-edge outboard sections. 

The roughness resulted in uniform separation such that there was no asymmetry. The effect on lift, shown 

in Figure 12 was similar to that observed by Koven and Graham. There was also a significant drag 

penalty for lift coefficients up to about 0.3 as indicated in Figure 13. The Reynolds number effects on 

maximum lift coefficient are summarized in Figure 5 for both the plain wing and with partial-span, 

inboard split flaps with and without roughness. The trend of increasing CL,max with Reynolds number for 

the clean wing is typical of what has been observed in numerous airfoil studies. The same is also true for 

the non-effect of Reynolds number for the wing with leading-edge roughness. It is also interesting that 
NASA/TM—2013-216381 18 

CL,max for the roughness leading edge with “flaps off” begins to decrease slightly for Re  8.0×106. Given 

observations of roughness effects on airfoils, this may be an effect of increasing Mach number 

(e.g., Ref. 2), which was not held constant in these experiments, but varied with Reynolds number. 

The early NACA experiments clearly show the general aerodynamic effect of roughness. While these 

wind-tunnel models may have been simplistic (for example, not employing wing twist) leading to tipstalling phenomena, the performance penalties due to roughness were not unrealistic. What is needed for 

the icing case is an understanding of the effects of roughness size, location, chordwise extent, 

concentration and possibly shape. Previous investigations have considered the roughness height of 

simulated upper surface wing frost. For example, Van Hengst et al., (Ref. 27) in their review article do 

summarize the effect of roughness size (k/cmac) on CL,max for various airplane configurations as shown in 

Figure 14. This plot is similar to the traditional “Brumby (Ref. 28) plot” that illustrates roughness size 

effects mostly on straight wings and airfoils. 
NASA/TM—2013-216381 19 

The effect of the chordwise extent/location of simulated frost roughness was investigated by Kind and 

Lawrysyn (Ref. 29) for a jet transport wing with an aspect ratio of 7.9, 30 sweep and taper ratio of 0.25. 

This was a numerical simulation, representative of wing frost with k/cmac = 0.00006 and the results are 

shown in Figure 15. The numbers in the figure legend indicate the chordwise extent of the roughness on 

the wing upper surface: “50” indicates the roughness extended from x/c = 0.42 to the trailing edge, “75” 

indicates x/c = 0.20 to the trailing edge, “95” indicates x/c = 0.009 to the trailing edge and “99” indicates 

x/c = 0.00 to the trailing edge. The data show the large sensitivity of CL,max and drag when the upstream 

edge of the roughness was located between x/c = 0.00 and 0.009. More data like the kind shown in 

Figures 14 and 15 are needed to better understand the effects of roughness size, location and 

concentration on swept-wing aerodynamics. 

Papadakis et al. (Ref. 30) performed some limited investigation of roughness size and chordwise 

extent of sandpaper roughness on the leading edge of a business jet T-tail configuration. The horizontal 

tail had a leading-edge sweep of 29, aspect ratio 4.4 and taper ratio of 0.43. Two models were tested. 

The full-scale model had a mean aerodynamic chord of 49.2 in. and the 25 percent scale model had a 

mean aerodynamic chord of 12.3 in. Both models had a fully-actuated elevator instrumented to measure 

hinge moments. Other data include forces, moments and surface pressures. Figure 16 compares the effect 

of chordwise extent of 40-grit sandpaper roughness (k/cmac = 0.000344) on lift coefficient for various 

Reynolds and Mach numbers. The Mach number varied proportionately with Reynolds number from 

M = 0.05 (at Re = 1.36×106

) to 0.23 (at Re = 6.60×106). The negative stall characteristics are shown in the 

plots as the critical characteristics for the horizontal tailplane. Note that the magnitudes of the stall angles 

and lift coefficients decreased as Reynolds number increased. Unfortunately, the authors provide no 

explanation for this effect that is different from that usually observed in airfoil roughness studies and the 

swept-wing studies cited in this report. Comparing the plots shows that there is little effect of chordwise 
NASA/TM—2013-216381 20 

roughness extent over the range tested. A much smaller roughness size of 120-grit sandpaper (k/cmac = 

0.000091) was tested in the same way. Comparing Figure 17 with Figure 16 shows that the smaller 

roughness did have a lesser effect on lift coefficient. In Figure 17, the lift coefficient at stall ranged from 

approximately –0.8 to –0.9 for the 120 grit roughness versus –0.7 to –0.8 for the 40 grit roughness with 

the same surface coverage in Figure 16. This is consistent with the data shown in Figure 14. 
NASA/TM—2013-216381 21 

While these roughness studies do provide some initial insights, more research is needed to understand 

the effect of size, location, concentration, Reynolds number and other factors important to the 

aerodynamics of iced-swept wings. There is very little or no flowfield information such as surface 

pressures, flow visualization and velocity profiles in these reports. An understanding of the roughness 

flowfield is important to accurate subscale simulation of roughness effects. 

3.2 Horn Ice 

On airfoils, glaze icing conditions often lead to leading-edge ice accretion known as horn ice. There are 

several geometric features that are important to the horn-ice shape, including height, the angle it makes with 

respect to the flow and its surface location (Ref. 2). For a swept wing in glaze or mixed icing conditions, ice 

accretion can be divided into three subcategories known as complete scallops, incomplete scallops and no 

scallops. This was discussed in Section 2.2 and illustrations of each shape are shown in Figure 8. The type 

that forms has been found to be dependent on sweep angle and icing conditions (Ref. 22). 

In contrast to ice roughness on swept wings, there is some fundamental research regarding the 

flowfield about a three-dimensional wing with horn ice. Bragg et al. (Ref. 31) performed a flowfield study 

on a swept wing using Laser Doppler Velocimetery (LDV) to measure all three components of velocity 

behind a simulated glaze ice shape. Their experimental results were complemented by the computational 

results of Kwon and Sankar (Ref. 32). The model tested was a rectangular semispan wing with a chord of 

15-in., sweep angle of 30 and an aspect ratio of approximately 2.3. A NACA 0012 airfoil section was 

used. The ice shape was produced by uniformly extending a horn ice shape formed on a NACA 0012 

airfoil in NASA’s Icing Research Tunnel. A cross section of the ice shape is shown in Figure 18. Since 

this ice shape was uniformly extended from a two-dimensional ice shape its cross-section did not change 

in the spanwise direction. 

A particle-trajectory simulation and surface-oil flow simulation from the CFD results are shown in 

Figure 19 at an angle of attack of 8. The flowfield over the iced-swept wing was characterized by a 

spanwise-running, leading-edge vortex formed from the separated shear layer initiated at the ice shape. 

The vortex was located progressively downstream moving outboard along the wing, eventually lifting off 

the wing surface. The spanwise location at which the vortex lifted off the surface moved inboard as the 

angle of attack increased. The simulated surface-oil flow showed regions of purely spanwise flow near the 

trailing edge and regions of almost completely reversed flow near the leading edge. These flowfield 

characteristics are analogous to the fundamental studies conducted by Poll (Ref. 18) and reviewed in 

Section 2.1 for the sharp radius leading-edge wing. There are also similarities to the separation bubble 

 
NASA/TM—2013-216381 22 

flowfield associated with horn ice on airfoils (Ref. 2). In the latter case, the lack of strong spanwise flow 

results in more significant differences at higher angles of attack approaching stall. 

Khodadoust and Bragg (Ref. 33) measured the pressure distribution at several spanwise locations as 

shown in Figure 20 for the same wing and horn ice used by Bragg et al. (Ref. 31) at the same conditions 

as the wing in Figure 19. The pressure distributions show that when the horn ice is present the leadingedge suction peaks are replaced by regions of nearly constant pressure covering approximately 20 to 

40 percent of the chord. These pressure plateaus indicate that a large vortex was present behind the ice 

shape which was corroborated by contour plots of the normalized streamwise velocity measured by LDV 

at y/b = 0.4 and y/b = 0.7 show in Figures 21 and 22, respectively. The acceleration of the flow over the 

ice shape can be seen near the leading edge at both spanwise locations. The recirculation region beneath 

the vortex is indicated by the negative streamwise velocity near the surface. A zero velocity contour can 

be found in the plots indicating the approximate streamline that divides the flow in the downstream 

direction and the recirculation region. The x/c location where this streamline intersects the model surface 

gives an idea of the vortex size. It can be seen that reattachment occurs at x/c = 0.23 and 0.77 for y/b = 0.4 

and 0.7, respectively. The increasing diameter of the vortex as the tip is approached is also seen in the 

CFD simulations and its influence is seen in the broadening and reduction of the pressure peak in 

Figure 20. Figures 19(a) and (b) show a very complex three-dimensional flowfield near the tip of the wing 

where the leading-edge vortex and tip vortex interact. LDV measurements in this area confirmed the 

interaction of these two vortices and measured maximum normalized spanwise velocities and turbulence 

intensities of approximately 0.62 and 35 percent, respectively. The individual pressure distributions and 

streamwise velocity contours at each spanwise locations were very similar in character to the large 

separation bubble formed behind horn ice on airfoils. The pressure plateau region is associated with the 

separated shear layer (upstream of transition) and the intersection of the zero-velocity contour with the 

surface is associated with the mean reattachment location (Ref. 2). Missing from the airfoil case is, of 

course, the spanwise velocity component that results in the coherent, spanwise-running, leading-edge 

vortex in the iced-swept-wing case. 

Papadakis et al. (Ref. 34) investigated the effect of horn ice shapes on a T-tail. The model was a 

25 percent scale business jet T-tail with a 29 leading-edge sweep, a mean aerodynamic chord of 

12.31 in. and an aspect ratio of 4.4. The ice shapes tested were generated using LEWICE (LEWICE 

(Ref. 35) is a computational tool that simulates ice growth on surfaces exposed to icing conditions), and 

in addition two spoilers consisting of a spanwise flat plate protruding normal to the airfoil surface were 

also tested. Cross-sectional views of the ice shapes and spoilers are shown in Figure 23. The ice-shape 

simulations were generated using the same atmospheric conditions but different simulation times. The 

spoilers were used to represent only the height of the horn on the suction surface. 
NASA/TM—2013-216381 23 

 
NASA/TM—2013-216381 24 

The effect of the ice simulations on lift at Re = 1.36×106 is shown in Figure 24. The ice shape 

designations L9 and L22 refer to LEWICE 9-min. and LEWICE 22-min. predictions; for the ice shapes 

designated L9B and L22B 24-grit (k/cmac = 0.0024) roughness was added. The negative stall 

characteristics are shown in the plots as the critical characteristics for the horizontal tailplane. The 

resulting performance degradation of the large ice shape (L22) is significantly worse than for the smaller 

shape (L9) and in both cases the roughness degrades performance even more at high angles of attack. 

Performance of the T-tail with the spoiler ice simulations (not shown here) was similar or worse than for 

the LEWICE shapes with roughness. The large effect of added roughness, not typically seen in past airfoil 

studies, may be due to the aerodynamics of a swept wing; however, another potential explanation could 

be the geometry of the LEWICE simulations. The horns of LEWICE shapes are significantly more 

rounded than those of typical ice accretion. Because the horns are fairly rounded the point of separation 

may not be fixed to a specific location on the horn. Therefore, the addition of roughness may affect the 

separation location, whereas on a sharper horn this point would be fixed regardless of the surface condition. 
NASA/TM—2013-216381 25 

 

Surface pressure data were acquired and are shown at the midspan station for α = –5 in Figure 25. 

Note the similarities to Figure 20 for the larger L22 and L22B configurations with the regions of 

approximately constant pressure on the upper surface being indicative of a leading-edge vortex. Pressure 

data for the smaller L9 and L9B configurations indicated a significantly smaller separated flow region. At 

this angle of attack, these pressure distributions do not exhibit large differences in performance between 

the smooth and rough configurations that were more obvious in the stall region. 

In addition to measuring the forces and surface pressures on the model, boundary-layer profiles were 

also measured. The boundary layer was measured at 60 percent chord at a spanwise location of 51 percent 

using a rake of total pressure probes. Boundary-layer profiles corresponding to the results in Figures 24 

and 25 at α = –5 are shown in Figure 26. It should first be noted that the boundary-layer velocities were 

calculated using the freestream static pressure and not the local static pressure so profiles are not true 

boundary-layer profiles; however, they still give insight into the thickness of the viscous region. Also, 

because measurements were made using total pressure probes, areas of reverse flow could not be 

measured therefore any zero readings above the surface indicate reverse flow. 

It can be seen in Figure 26 that the boundary layers for the iced cases are significantly thicker than the 

clean case. When comparing the boundary-layer profile between the two different ice shapes, the 22-min. 

ice shape (L22) still had regions of reverse flow, indicated by the areas of zero measured velocity above 

the surface. This means that the leading-edge vortex has not reattached by x/c = 0.6 for the large ice 

shape, whereas it has reattached upstream of x/c = 0.6 for the smaller ice shape (L9) at this spanwise 

station. These data, combined with the pressure distributions in Figure 25, show trends similar to that 

described in connection with Figures 20 to 22. For both ice shapes the boundary-layer thickness of the 

rough ice was slightly larger than that of the smooth ice; however, it is unclear if the measured boundarylayer profiles would be more similar if based upon local static pressure as opposed to the freestream static 

pressure. 

 
NASA/TM—2013-216381 26 

The previously cited works of Bragg et al. (Refs. 31 and 33) and Papadakis et al., (Ref. 34) utilized 

horn ice shapes that were essentially two-dimensional, similar to glaze ice accretion with no scallops. 

Furthermore, the results were analogous to those in the large existing body of work for horn ice on 

airfoils. A later study by Papadakis et al. (Refs. 36 and 37) investigated the effects of several different ice 

shapes on a swept wing that included high-fidelity, three-dimensional ice-casting simulations 

incorporating various levels of incomplete and complete scallops. The swept-wing model had a GLC-305 

airfoil section aligned in the streamwise direction with a 28 leading-edge sweep, 18.72-in. mean 

aerodynamic chord, 60-in. semispan and an aspect ratio of 6.8. The ice shapes used for this study were 

formed on the same model in the NASA Glenn Icing Research Tunnel. The icing conditions were chosen 

to produce a wide range of ice shapes including complete and incomplete scallops of various sizes and a 

rime ice formation. Moldings of the IRT ice accretions were subsequently used to produce the ice castings 

that were used for the aerodynamic tests. The ice-casting simulations captured the fully three-dimensional 

variation of the ice accretion. 

The ice accretions tested and the corresponding tunnel conditions are shown in Table 1. Ice1 (IRTCS10) is an example of a complete scallop glaze ice accretion. This shape can be compared with Ice2 

(IRT-IS10) which is an example of an incomplete scallop. While the cross-sectional tracings look similar, 

the photographs reveal the large differences in the three-dimensional geometry. Ice3 (IRT-SC5) is a rime 

ice accretion and will be discussed in Section 3.3. Ice4 (IRT-CS2) was formed at identical conditions to 

Ice1(IRT-CS10), but with a much shorter exposure time of 2 min. Therefore, this is a relatively small 

glaze ice accretion and it is unclear if true scallop features formed due to the short exposure time. Ice5 

(IRT-CS22) was also formed at identical conditions to Ice1 (IRT-CS10), but with a longer exposure time 

of 22.5 min. This longer duration led to the very large complete scallop accretion shown in Table 1. 

Finally, Ice6 (IRT-IPSF22) had some characteristics of complete scallops that were not as fully developed 

as in Ice1(IRT-CS10) and Ice5 (IRT-CS22). 

In addition to the IRT ice accretion in Table 1, seven additional ice shapes were generated using the 

ice accretion code LEWICE 2.0. Since LEWICE 2.0 is a two-dimensional ice accretion code, the threedimensional simulations used for the aerodynamic testing were composed of two-dimensional slices at 

several spanwise sections of the wing blended together. The detailed procedure is discussed in the original 
NASA/TM—2013-216381 27 

TABLE 1.—ICE ACCRETIONS AND CORRESPONDING ICING CONDITIONS USED BY PAPADAKIS ET AL. (REF. 37). 

reference (Ref. 36). The LEWICE shapes were formed at the same conditions shown in Table 1, but the 

velocity and airfoil geometry used in the computations was the velocity normal to the leading edge. The 

seventh ice accretion used the same conditions as Ice Shape 5 (IRT-CS22) in Table 1 but the streamwise 

velocity was used. The LEWICE shapes were essentially horn ice simulations with no scallop features. 

These are designated with the prefix “LS” in the following figures. The geometry did vary in the spanwise 

direction but only according to the number of slices used to loft the three-dimensional geometry. The 

effect of ice-shape roughness was also investigated by adding 36-grit size roughness (k/cmac = 0.0011) to 

the smooth LEWICE shapes. These are designated with the prefix “LR” in the following figures. 

Because the Papadakis et al. (Refs. 36 and 37) study utilized high-fidelity ice-casting simulations of 

the three-dimensional ice accretions, it is possible to make some observations about the attendant 

aerodynamic effects relative to horn ice with no scallops or minimal three-dimensional features. Figure 27 

shows a comparison of the ice accretion cross-section and the corresponding LEWICE shape for Ice2 

(IRT-IS10). Overall, there is reasonable agreement in the tracings, particularly for station A (50 in. from 

root). A comparison based upon aerodynamic effect is shown in Figure 28 for lift, drag and pitchingmoment coefficients. The results here are very similar to what has been reported for horn ice simulations 

on airfoils (Ref. 9). The effect on lift coefficient is very similar for all three configurations up to the stall 

region where the three-dimensional ice casting configuration has slightly lower maximum lift and stalling 
NASA/TM—2013-216381 28 

angle. For horn ice shapes on airfoils, small differences in the horn height, location or angle can 

significantly affect CL,max and may explain the differences given the obvious differences in the cross 

sections shown in Figure 27. The smooth LEWICE (LS-IS10) shape had slightly higher CL,max than the 

LEWICE shape with roughness (LR-IS10), which is consistent with the data shown in Figure 24 for the 

business jet T-tail and for past research on airfoils. For the drag, the three-dimensional ice casting 

configuration had slightly higher drag coefficient at low CL where differences in profile drag are more 

pronounced. This trend was typically observed for horn ice simulations on airfoils with more threedimensional simulations having higher profile drag, especially at low lift coefficient. The pitching 

moment data reflects the most significant differences among the three configurations that are indicative of 

the differences in the distributions of surface pressure. 

 
NASA/TM—2013-216381 29 

Also included in Figure 28 is a comparison of surface pressure at the 50 percent semispan stations at 

α = 4. For all three configurations, there are fairly well defined regions of approximately constant 

pressure from x/c ≈ 0.0 to x/c ≈ 0.2 on the upper surface. This pressure signature is similar to that shown 

previously in Figures 20 and 25 and is due to the leading-edge vortex induced by the horn geometry. In 

their report, Papadakis et al. (Ref. 37) provide more pressure comparisons at several angles of attack and 

for two additional spanwise locations. In most of these plots, the general shape of the distributions for the 

Ice2 (IRT-IS10) casting is similar to that for the LEWICE simulations. Thus indicating that for 

incomplete scallops, the large flowfield features (such as the leading-edge vortex) are similar to the horn 

with no scallops described earlier in this section. There were significant differences in the magnitudes of 

the pressure coefficient as evidenced by the pitching-moment data shown in Figure 28. It is the general 

topologies of surface pressure that are similar. 

 
NASA/TM—2013-216381 30 

The data presented in Figures 27 and 28 for the incomplete scallop configuration of Ice2 (IRT-IS10) 

are contrasted with the data for the complete scallop configuration of Ice1 (IRT-CS10). The cross-section 

geometry and LEWICE results are depicted in Figure 29 where there are much larger differences in the 

two geometries relative to that shown in Figure 27. This is most likely a result of the much higher degree 

of three-dimensionality associated with the complete scallop condition. It is not expected that the 

LEWICE cross sections in Figure 29 should match since LEWICE does not generate three-dimensional 

results. Even the most sophisticated ice accretion simulation codes do not predict the level of threedimensionality associated with compete scallop ice accretion. It is also very difficult to compare a twodimensional tracing of a highly three-dimensional ice accretion as small changes in spanwise location can 

have a very large effect on the cross-sectional tracing. 

 
NASA/TM—2013-216381 31 

Given these differences in geometry, it is interesting to compare the aerodynamic results in Figure 30. 

For example, the main effect on lift coefficient is not observed until the stall region where the differences 

in CL,max among the three configurations is larger than for the incomplete scallop case in Figure 28. This 

difference is also observed in the drag coefficient in the stall region and is most likely due to the contribution 

of the induced drag. The pitching-moment variation reveal more significant differences among the 

configurations which are expected given the differences in the ice-shape geometry shown in Figure 29. 

 
NASA/TM—2013-216381 32 

The selected pressure distribution shown in Figure 30 is typical of the others plotted by Papadakis 

et al. (Ref. 37) In this case, there is a sharp contrast between the pressure signatures of the Ice1 (IRTCS10) casting and the LEWICE ice-shape simulations. The upper-surface pressure distribution of the 

LEWICE configurations are very similar to what has been shown previously in Figures 20, 25, and 28. 

This is particularly true for the case with roughness, LR-CS10. But for the Ice1 (IRT-CS10) casting, there 

is a gradual pressure recovery region with peak values of suction pressure much less than for the 

LEWICE shapes. A similar comparison is shown in Figure 31 which is for Ice6 (IRT-IPSF22). This 

condition also had a highly three-dimensional ice accretion geometry typical of complete scallops. In this 

case, both the smooth and rough LEWICE shapes (LS-IPSF22 and LR-IPSF22) exhibited a clearly 

identifiable region of approximately constant pressure from x/c ≈ 0.06 to x/c ≈ 0.2 associated with the 

leading-edge vortex. In contrast, the pressure distribution for the Ice6 (IRT-IPSF22) casting simulation 

was more similar to that shown in Figure 30 for the complete scallop case. Since there are significant 

differences in cross-section geometries between the ice castings and the LEWICE simulations, better 

matching of the surface pressure distributions is not expected. What is of interest here is that the shape of 

the pressure distributions is different in the area immediately downstream of the ice shape. For the 

nominally two-dimensional LEWICE shapes, there is a fairly distinct region of approximately constant 

pressure that has been shown to correspond to the leading-edge vortex. For the highly three-dimensional 

geometry of the complete scallop ice casting simulations, this pressure region was not observed and the 

peak suction pressures were much lower. To the authors’ knowledge, the flowfield immediately 

downstream of the highly three-dimensional geometry has not been characterized or reported in the 

technical literature. These pressure data suggest that there exists a significantly different flowfield than 

for the nominally two-dimensional horn ice shapes reported in past swept-wing icing studies. More 

research is needed to understand the flowfield associated with highly three-dimensional horn ice on swept 

wings so that the resulting impacts on wing performance may be better explained. 

 
NASA/TM—2013-216381 33 

Since current state-of-the-art ice accretion simulation codes are unable to predict highly threedimensional features such as complete scallops, it is important to understand their impact on the 

aerodynamics. The authors are unaware of any comprehensive study in this regard. However, Papadakis 

et al. (Refs. 36 and 37) did conduct a limited study on the effects of the gap between scallop tips by 

progressively filling in the gaps at various spanwise locations with a modeling compound to form a solid 

horn ice shape similar to those formed on airfoils. Measurements were made using Ice5 (IRT-CS22, 

Table 1), with the mid-section gaps closed, then with mid- and tip-section gaps closed, and then with all 

of the gaps closed. Progressively filling the gaps did result in minor increases in lift coefficient, but minor 

increases in drag coefficient were also observed. The Ice5 (IRT-CS22) configuration is a very large ice 

shape, so it is difficult to determine the significance of these results. 

In terms of surface pressure distribution, Papadakis et al. (Refs. 36 and 37) observed similar effects 

shown in the pressure distributions of Figures 30 and 31. The pressure signatures with the gaps filled 

tended to look like the LEWICE shape pressure signatures shown in Figures 30 and 31. The authors 

explain that the lower suction pressures immediately downstream of the ice shape associated with the 

highly three-dimensional scallop shapes, were due to leakage of air from region of high pressure, at the 

forward face of the ice shape, to low pressure downstream. Since the gap-filled shape, or LEWICE 

shapes, were “solid” or more nominally two-dimensional, this leakage did not occur leading to the higher 

suction pressures immediately downstream of the ice shape. Papadakis et al. (Refs. 36 and 37) suggest 

that this may be responsible for the higher values of lift coefficient observed for the nominally twodimensional configurations. Given the large size of this ice shape relative to the wing chord the observed 

effects and corresponding explanation may not be similar in other cases. Clearly, more research is needed 

to better understand the effect of complete scallops on iced-wing performance. 

Unlike the situation for other ice accretion on swept wings, there are currently both experimental 

performance measurements as well as flowfield studies for horn ice; although these studies offer valuable 

insights into the effect of horn ice accretions on swept wings there is still a substantial lack of information 

relative to the airfoil case. For example, a shortcoming of the studies discussed in this section is that no 

attempt was made to ensure that the IRT generated ice shapes were accurate representations of full-scale 

ice accretion. This could impact the aerodynamic effects of complete scallops. Other areas where 

experimental data are needed include more flowfield studies for detailed ice shapes with and without 

scallops and different wing geometries, parametric studies investigating geometrical features of the ice 

such as height, shape (e.g., tip radius, roughness level), location and the influence of wing geometry. 

The data presented in this section for horn ice on swept wings suggest an aerodynamic 

subclassification that distinguishes between “nominally three-dimensional” horn ice and “highly threedimensional” horn ice. Nominally three-dimensional horn ice is associated with glaze ice accretion having 

either no scallops or incomplete scallops. This is only nominally three-dimensional since the gross shape 

does not vary significantly over small spanwise distances. Highly three-dimensional horn ice is associated 

with glaze ice accretion having complete scallop tips. In this case, the ice geometry changes significantly 

over small spanwise distances. From the perspective of aerodynamic classification, the difference 

between nominally three-dimensional and highly three-dimensional horn ice is defined in terms of the 

flowfield characteristics. The results discussed in this section for Figures 18 to 28 were all associated with 

nominally three-dimensional horn ice as there was very little change in the simulated ice geometry in the 

spanwise direction. The flowfield was described in terms of the leading-edge vortex and the resulting 

region of approximately constant surface pressure aft of the ice shape. This was contrasted against some 

results for highly three-dimensional horn ice in Figures 29 to 31 where the pressure signatures exhibited 

more gradual pressure recovery in place of the constant pressure regions aft of the ice shape. Due to a lack 

of data for highly three-dimensional horn ice on a swept wing, the key flowfield characteristics in this 

case are unknown. To further distinguish between nominally three-dimensional and highly threedimensional horn ice flowfield characteristics is yet another area where further research is required. 
NASA/TM—2013-216381 34 

3.3 Streamwise Ice 

Streamwise ice is most often associated with rime icing conditions. Streamwise ice typically follows 

the wing leading edge contour or forms a horn-like shape, or protuberance, oriented into the flow 

direction. As discussed in Section 2.2 and illustrated in Figures 6 and 7, the region of the attachment line 

for swept wings may not be smooth as is usually the case for streamwise ice on airfoils. Streamwise ice 

does not have scallops, gaps, or other highly three-dimensional variations in the geometry along the span. 

Streamwise ice on swept wings has received very little attention in the literature. As in the case of 

airfoils, it can be anticipated that the aerodynamic effects of streamwise ice will be small relative to horn 

ice. The study by Papadakis et al. (Refs. 36 and 37) included one streamwise ice shape, Ice3 (IRT-SC5) in 

Table 1. The aerodynamic effects of Ice3 (IRT-SC5) as well as the smooth and rough LEWICE 

simulations (LS-SC5 and LR-SC5) are shown in Figure 32. For this case, there was very good agreement 
NASA/TM—2013-216381 35 

in the cross-sections of the IRT ice accretion and the LEWICE simulations as shown in Figure 32. The lift 

data in Figure 33 show that that the streamwise ice actually increased the maximum lift coefficient and 

stalling angle of the wing. Relative to the clean wing the IRT ice-casting simulations increased CL,max by 

3.4 percent and the smooth and rough LEWICE simulations increased CL,max by 8.0 and 12.6 percent, 

respectively. An increase in CL,max due to streamwise ice has also been observed on airfoils (Ref. 38) and 

has been attributed to the ice shape effectively forming a leading-edge flap with the increase in chord 

length and wing area relative to the reference area. Another potential factor in this case is the lowReynolds number (1.8×106

) at which the performance data were acquired. It is expected that the 

maximum lift and stalling angle of the clean wing would increase significantly at higher Reynolds 

number. Some change in the iced wing may be expected, but may not be as large given the Reynolds 

number effects for leading-edge roughness described in Section 3.1. All three ice shapes increased drag at 

the lower angles of attack, but for angles of attack greater than approximately 10, the rough LEWICE 

shape had nearly the same drag as the clean wing and the smooth shape actually decreased the drag. 
NASA/TM—2013-216381 36 

Pressure distributions for the three streamwise ice shapes at 50 percent semispan are shown in 

Figure 34. At α = 4 and 8 the pressure distributions of the clean and iced wings are in good agreement at 

all spanwise locations except near the leading edge where the minimum pressure of the clean wing is less 

than the iced wing. Papadakis et al. (Ref. 36) note that this is due to a decrease in the effective angle of 

attack resulting from the leading-edge droop caused by the ice shape. At α = 12 it appears that a small 

leading-edge vortex is present for both the clean and iced wings although greater suction is present on the 

iced wings. The suction is greatest for the smooth LEWICE shape. Although not shown here, at 

15 percent semispan only, the IRT ice-casting simulation produced a leading-edge vortex, with 

reattachment located progressively downstream from root to tip. It should also be noted that the three ice 

shapes showed greater pressure recovery at the trailing edge. This was also true at the 15 percent 

semispan station but could not be determined at the 85 percent semispan station because the pressure taps 

ended at x/c = 0.6 due to the aileron. Finally, at α = 16 the clean wing was completely separated at the 

50 percent semispan location but the iced wing flow remained at least partially attached. At the 85 percent 

semispan location all conditions were completely separated over at least the first 60 chord. 
NASA/TM—2013-216381 37 

The results shown here clearly illustrate the need for more aerodynamic data for streamwise ice on 

swept wings. It is expected that streamwise ice will not improve maximum lift performance in most cases. 

More information about the flowfield is required to understand the effects of roughness on streamwise ice 

that are thought to be important to the aerodynamics. For streamwise ice on airfoils, the aerodynamic 

effects were chiefly made manifest through trailing-edge separations. While small leading-edge 

separation bubbles were often observed at the ice/airfoil juncture this flowfield feature did not play a 

decisive role in the resulting aerodynamics. For swept wings, it is expected that this separation, if present, 

may lead to leading-edge vortex formation. It remains to be determined which of these effects, or perhaps 

both, contribute to the observed performance changes on the iced swept wing. 

3.4 Spanwise-Ridge Ice 

Spanwise-ridge ice can be associated with a number of icing conditions where the wing leading edge 

is free of ice with sometimes large ice formations located farther downstream. Typical examples are SLD 

icing conditions coupled with ice-protection system operation. Large drops can impinge on the wing aft of 

the protected areas sometimes forming an ice accretion best described as a ridge. Spanwise-ridge ice can 

also form when a heated leading-edge, ice-protected surface is not evaporating all of the impinging water. 

The liquid water flows downstream from the ice-protection system where it freezes forming a ridge 

oriented in the spanwise direction. Spanwise-ridge ice accretions have been extensively studied on airfoils 

(Refs. 14, 39, and 40). The spanwise-ridge ice flowfield was investigated by Lee et al., (Ref. 41) and it 

was shown that the boundary layer separates from the ice accretion and a separation bubble is formed 

downstream of the ice accretion. It was observed that the size of the separation bubble grew as angle of 

attack increased, similar to the separation bubble downstream of horn ice. Many of the investigations 

have involved parametric studies of the effects of ridge height, location and airfoil geometry. A key 

finding in the airfoil research was that the location of the ridge relative to clean pressure loading indicated 

the severity of the lift degradation due to the ridge. Airfoils designed with “forward loading” were much 

more sensitive to spanwise ridges located in the first 10 to 15 percent chord relative to airfoils that had a 

more uniform or aft loading. Johnson (Ref. 42) and Lee et al. (Ref. 43) both investigated the effects of 

spanwise-ridge ice accretions on straight wings. Both studies found large decreases in lift, however no 

parametric study investigating the influence of various ice-shape features was performed. Since the wings 

incorporated little or no sweep, the basic aerodynamics was likely similar to that for airfoils. There is very 

little information regarding the effects of spanwise-ridge ice accretions on swept wings, and the few 

studies that have been performed only focused on the aerodynamic performance of the wing with no 

corresponding flowfield investigation. 

Papadakis et al. (Ref. 44) performed a parametric study of spanwise-ridge ice on the same, swept 

GLC-305 model used by Papadakis et al. (Ref. 36). Ice shapes were simulated by uniformly extending the 

simple geometries, shown in Table 2, across the span of the wing. The method of simulating spanwiseridge ice with simple geometries is common practice for airfoils (Ref. 40). It is unknown if any effort was 

made to use simulations that accurately represented any documented full-scale ice accretion. The ridge 

heights ranged from 0.2 to 0.5-in. corresponding to 1 to 2.7 percent of the 18.72-in. mean aerodynamic 

chord. Each ice-shape simulation was tested at 2.5, 5, 10, 15, 20, and 30 percent chord measured in the 

streamwise direction. All of the ice shapes tested resulted in reduced lift curve slopes and increased drag 

regardless of the chordwise location; however, for most of the ice shapes and locations tested, in stark 

contrast to results presented by Johnson (Ref. 42 and Lee Ref. 43) for straight wings, the ice shapes 

increased the maximum lift coefficient relative to the clean wing. The effects of the simulated ice shapes, 

located at 15 percent chord, on lift are shown in Figure 35. It can be seen that although the lift curve slope 

was reduced for the iced cases, the stalling angle of attack and maximum lift coefficient increased. For all 

angles of attack below clean wing stall (13.8) the drag of the iced wing was significantly higher than the 

clean wing. It is also interesting that for all except the largest ice shape (RB-6), the pitching moment 

variation with angle of attack was mostly unaffected up to stall. 
NASA/TM—2013-216381 38 

TABLE 2.—SIMULATIONS OF SPANWISE RIDGE ICE 

USED BY PAPADAKIS (REF. 44). 

Increases in lift coefficient at high angles of attack have been observed for streamwise ice accretions 

on airfoils (Ref. 38) and a wing (Section 3.3) (Ref. 36). It has also been observed for spanwise-ridge ice 

accretions on airfoils. Whalen (Ref. 40) and Papadakis et al. (Ref. 41) investigated the effects of 

simulated spanwise-ridge ice accretions on airfoils at Reynolds numbers of 1.8×106 and 2.0×106

, 

respectively. They both observed that when the clean airfoil exhibited trailing-edge stall and the height of 

the ice simulation was comparable to the local boundary-layer thickness, the iced airfoil performed better 

at angles of attack near and above clean wing stall. This performance enhancement was attributed to the 

mixing layer generated by the ice shape entraining higher-momentum fluid into the boundary layer. This 

explanation may not suffice for the case of the ice simulations shown in Table 2. The size of these 

simulations ranged from 0.2 to 0.5 in., likely making them significantly larger than the local boundary 

layer. It is also important to note that Broeren et al. (Ref. 14) showed that this lift enhancing effect on 

airfoils can be an artifact of low-Reynolds number testing of the clean airfoil. Their results showed that it 

is possible for the iced airfoil to have better high angle of attack performance characteristics than the 

clean airfoil at a Reynolds number of 1.8×106

, but when compared to the clean airfoil at a Reynolds 

number of 15.9×106

 the iced airfoil performance degraded substantially. Due to these results it is difficult 

to determine the exact cause of the performance enhancement seen by Papadakis et al. (Ref. 44). This also 

emphasizes the importance of investigating Reynolds number effects where such data are lacking. 

Despite the problems interpreting the integrated performance characteristics of the wing with 

simulated ice shapes, the pressure distributions are qualitatively similar to the pressure distributions of 

airfoils with spanwise-ridge ice accretions which offers valuable insights into the flowfield. Figure 36 

shows the chordwise pressure distribution at 15 percent semispan for the clean wing and each of the ice 

simulations placed at 15 percent chord and at an angle of attack of 4. There are several features worth 

pointing out when the ice shape is present. First, the upper surface pressure in the region upstream of the 

ice simulation located at 15 percent chord is significantly greater than the pressure on the clean wing. This 

is due to the flow deceleration as it approaches the ice accretion. This is immediately followed by a large 

drop in pressure resulting from the acceleration of the flow over the ice simulation. These pressure 

distributions are very similar to pressure distributions seen on “short” spanwise-ridges as described by 
NASA/TM—2013-216381 39 

Broeren et al. (Ref. 14) on airfoils. For several of the ice simulations, most noticeably for the RB-6, there 

is a region of approximately constant pressure on the upper surface followed by rapid pressure recovery. 

The pressure coefficient approaches the clean wing values near the trailing edge. This pressure signature 

is analogous to that associated with the leading-edge vortex described extensively in Section 3.2 for horn 

ice. However, the authors are unaware of any existing flowfield data that show the presence of such a 

vortex flow for spanwise-ridge ice on swept wings. 
NASA/TM—2013-216381 40 

There are currently several areas where there is either no data or insufficient data regarding spanwiseridge ice shapes. The parametric studies by Papadakis et al. (Ref. 44) must be interpreted cautiously due 

to the low-Reynolds number and counter-intuitive results that contradict equivalent airfoil results with no 

flowfield explanation. Future research should explore Reynolds number effects. The authors are unaware 

of any swept-wing studies that used high-fidelity ice shapes such as those formed from castings of actual 

ice accretions. As a result, there are no data that can be used to validate the method of simulating ice 

shapes with simple geometries on swept wings. It is well known that real spanwise-ridge type ice 

accretion can be highly three-dimensional with significant spanwise variation in the gross shape. Further 

parametric studies are needed to determine the effects of ridge ice height, shape and wing geometry. Such 

work must be supplemented with extensive flowfield studies in order to improve our understanding of 

how the ice accretions affect the aerodynamics of swept wings. Of particular interest is the spanwise 

vortex interaction with the spanwise ridge in determining the attendant aerodynamic effects. 

4.0 Summary 

The continued design, certification and safe operation of swept-wing airplanes in icing conditions rely 

on the advancement of computational and experimental simulation methods for higher fidelity results 

over an increasing range of aircraft configurations and performance, and icing conditions. There is 

increasing demand to balance trades-offs in aircraft efficiency, cost and noise that tend to compete 

directly with allowable performance degradations over an increasing range of icing conditions. Aircraft 

icing research has now reached the level of maturity that computational methods and experimental tools 

are currently being used to address many of these challenges. However, knowledge gaps do remain for 

swept-wing geometries and larger droplet icing conditions. The current state-of-the-art in icing 
NASA/TM—2013-216381 41 

aerodynamics is mainly built upon a comprehensive understanding of two-dimensional geometries 

developed from myriads of research efforts described in the technical literature. Such an understanding 

for fundamentally three-dimensional geometries such as swept wings does not currently exist. The 

purpose of this report is to describe what is known of iced-swept-wing aerodynamics; to identify the type 

of research that is required to improve the current understanding; and to develop an aerodynamically 

based classification of swept-wing ice accretion. This report focuses on the fundamental aerodynamics of 

iced swept wings. The existing data tend to be: (1) mostly at low-Reynolds number and (2) applicable to 

simple swept-wing geometries that do not have high-lift systems, wing-mounted engines, fuselages and 

other features of actual airplane wings. These factors can significantly alter the iced aerodynamics for 

particular configurations and so extreme caution must be exercised in terms of making general 

conclusions based upon the current, limited database. 

This report also describes two additional aspects to swept-wing aerodynamics: (1) the fundamental 

stalling mechanisms and flowfield features of clean swept wings and (2) the aerodynamic characteristics 

of swept-wing ice accretion. The NACA and other organizations carried out fundamental research into the 

stalling characteristics of swept wings. The chief stalling mechanisms have been divided into leadingedge and trailing-edge separations. A key characteristic of leading-edge separations is the spanwiserunning, leading-edge vortex that lies close to the wing surface prior to stall. Leading-edge separations 

often result from sharp or small leading-edge radii and there is some evidence of the Reynolds number 

independence of these flowfields. This reduced Reynolds number effect may also be analogous to iced 

swept wings where the separation point is often set by the ice geometry. This report details some of these 

comparisons. 

Ice accretion formations on swept wings can have unique characteristics. Depending upon specific 

icing conditions and sweep angle, the region of the attachment line may not be smooth as is often the case 

for airfoils. While initial roughness and rime ice accretion on swept wings tend to look very similar to that 

on airfoils, there can be significant differences for glaze ice accretion. For glaze icing, certain 

combinations of icing conditions and sweep angle can lead to the formation of highly three-dimensional 

features called scallops that do not exist for ice accretions on airfoils. It is also possible to have glaze ice 

accretion with no scallops or even incomplete scallop formations on swept wings. 

Following the method used in a previous review of iced-airfoil aerodynamics, this report classifies 

swept-wing ice accretion into four groups that are based upon unique flowfield features. Instead of relying 

upon ice accretion terminology such as rime and glaze, the four aerodynamic groups have names 

associated with ice-shape geometry. These four groups are: ice roughness, horn ice, streamwise ice and 

spanwise-ridge ice. This report attempts to describe the unique flowfield features of each group that 

determines the iced-wing aerodynamics: 

 Ice roughness represents initial leading-edge ice accretion and a key aerodynamic characteristic is that 

the scale of the boundary-layer separation is of the same order as the size of the roughness. While 

there are many studies that have looked at roughness effects on swept-wing performance, including 

Reynolds number effects, there is a lack of flowfield data from which to interpret these results. More 

data are needed to understand the effects of roughness size, location and concentration on swept-wing 

aerodynamics. 

 Horn ice is large, leading-edge ice accretion that can be associated with glaze icing conditions. The 

flowfield is characterized by large-scale, boundary-layer separation originating at the horn. This 

separation leads to the formation of a spanwise-running, leading-edge vortex that is similar to that 

found on clean swept wings with leading-edge separations. There are a number of low-Reynolds 

number studies that have characterized the horn-ice flowfield for swept wings and documented the 

behavior of the leading-edge vortex preceding wing stall. This presents an excellent starting point, 

especially for nominally three-dimensional horn shapes such as those with no scallops or even 

incomplete scallops. However, there are no flowfield data known to the authors for highly threedimensional horn ice such as complete scallop formations. Therefore, the fundamental aerodynamics 

are essentially unknown in this case. This is an important factor since the associated performance 
NASA/TM—2013-216381 42 

penalties may be large. The small amount of existing data indicate that there are fundamental 

flowfield differences between nominally three-dimensional horn ice characterized by no scallop 

formations versus highly three-dimensional horn ice characterized by fully developed scallop 

formations. Flowfield data, such as mean and fluctuating velocity profiles and surface shear stress are 

needed to further understand the important differences observed between these two cases of horn ice 

on swept wings. 

 Streamwise ice can be associated with rime icing conditions and is generally conformal to the wing 

leading edge, or may form a horn-like feature (or protuberance) oriented into the flow direction. The 

only example of this group cited in this report showed an increase in wing maximum lift coefficient 

with the streamwise ice. While this effect may be possible, it is not expected to hold for most cases 

and illustrates the need for further wing performance data and flowfield information with realistic 

streamwise ice simulations. 

 Spanwise-ridge ice can be associated with ice protection system operation in SLD icing conditions or 

incomplete evaporation of impinging water. The leading edge is free of ice with an ice ridge located 

downstream often in the range of 10 to 15 percent chord. This report describes data from only one 

low-Reynolds number study for very simple geometric representations of spanwise-ridge ice on a 

swept wing. More aerodynamic performance data and flowfield information are needed for realistic 

spanwise-ridge ice simulations. 

For all of the proposed ice-shape classifications, relatively little is known about the three-dimensional 

flowfield and even less about the effect of Reynolds number and Mach number on these flowfields. Both 

of these deficiencies are important and limit the ability to classify swept-wing ice accretion. Most of the 

data found in the literature pertain only to aerodynamic performance. Except for nominally threedimensional horn ice, flowfield information is limited to some pressure distributions, all at low-Reynolds 

number. Variations in Reynolds number found for iced-swept wings are all for relatively low-Reynolds 

number and provide no guidance as to the appropriateness of these data at Reynolds numbers approaching 

flight. In the two-dimensional case, Reynolds and Mach number effects have been shown to be small in 

most cases and low-Reynolds number data have been used extensively to classify ice shapes and improve 

our understanding of iced-airfoil flowfields and aerodynamics. The very limited data available on swept 

wings to date suggest a similar result, but much more data are needed, particularly for realistic ice-shape 

simulations at higher Reynolds numbers. 

The classifications and supporting data presented in this report can serve as a starting point as new 

research explores swept-wing aerodynamics with ice shapes. As further results become available, it is 

expected that these classifications will need to be revised just as has occurred in the airfoil case. 

References 

1. Lynch, F.T., and Khodadoust, A., “Effects of Ice Accretions on Aircraft Aerodynamics,” Progress in 

Aerospace Sciences, Vol. 37, No. 8, Nov. 2001, pp. 669-767. 

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

Aerospace Sciences, Vol. 41, No. 5, Jul. 2005, pp. 323-362. 

3. Bragg, M.B., Broeren, A.P., and Blumenthal, L.A., “Iced-Airfoil and Wing Aerodynamics,” 

SAE Paper 2003–01–2098, Jun. 2003. 

4. Bragg, M.B., and Gregorek, G.M., “Environmentally Induced Surface Roughness Effects on Laminar 

Flow Airfoils: Implications for Flight Safety,” AIAA Paper 89–2049, Jul.-Aug. 1989. 

5. Papadakis, M., Yeong, H.W., Chandrasekharan, R., Hinson, M., Ratvasky, T.P., and Giriunas, J., 

“Experimental Investigation of Simulated Ice Accretions on a Full-Scale T-tail,” AIAA Paper 2001-

0090, Jan. 2001. 

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

With Simulated Ice Accretions,” AIAA Paper 2001-0833, Jan. 2001. 
NASA/TM—2013-216381 43 

7. Papadakis, M., Alansatan, S., and Yeong, H.W., “Aerodynamic Performance of a T-tail With 

Simulated Ice Accretions,” AIAA Paper 2000-0363, Jan. 2000. 

8. 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, Jan. 2007. 

9. 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. 

10. 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 2001-0727, Jan. 2003, also NASA/TM—2003-212124, Feb. 

2003. 

11. Bragg, M.B., and Khodadoust, A., “Experimental Measurements in a Large Separation Bubble Due to 

a Simulated Glaze Ice Shape,” AIAA Paper 88-0116, Jan. 1988. 

12. Broeren, A.P., Bragg, M.B., and Addy, H.E., Jr., “Flowfield Measurements About an Airfoil With 

Leading-Edge Ice Shapes,” Journal of Aircraft, Vol. 43, No. 4, Jul.-Aug. 2006, pp. 1226-1234. 

13. Lee, S., and Bragg, M.B., “Investigation of Factors Affecting Iced-Airfoil Aerodynamics,” Journal of 

Aircraft, Vol. 40, No. 3, May-Jun. 2003, pp. 499-508. 

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

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

15. 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. 

16. 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. 

17. Furlong, G.C., and McHugh, J.G., “A Summary and Analysis of the Low-Speed Longitudinal 

Characteristics of Swept Wings at High Reynolds Number,” NACA TR-1339, 1952. 

18. Poll, D.I.A., “Spiral Vortex Flow Over a Swept-Back Wing,” Aeronautical Journal, May 1986. 

19. Broeren, A.P., Lee, S., Shah, G.H., and Murphy, P.C., “Aerodynamic Effects of Simulated Ice 

Accretion on a Generic Transport Model,” SAE Paper 2011-38-0065, June 2011. 

20. Polhamus, E.C., “A Survey of Reynolds Number and Wing Geometry Effects on Lift Characteristics 

in the Low-Speed Stall Region,” NASA CR-4745, June 1996. 

21. Neely, R.H, and Connor, D.W., “Aerodynamic Characteristics of a 42 Swept-Back Wing With 

Aspect Ratio 4 and NACA 641-112 Airfoil Sections at Reynolds Numbers From 1,700,000 to 

9,500,000,” NACA RM L7D14, 1947. 

22. Vargas, M., “Current Experimental Basis for Modeling Ice Accretions on Swept Wings,” Journal of 

Aircraft, Vol. 44, No. 1, Jan.-Feb. 2007, pp. 274-290. 

23. Vargas, M., Papadakis, M., Potapczuk, M.G., Addy, H.E. Jr., Sheldon, D., and Giriunas, J., “Ice 

Accretions on a Swept GLC-305 Airfoil.” SAE 02GAA-43, SAE General Aviation Technology 

Conference and Exhibition, Wichita, KS, Apr. 2002, Also NASA/TM—2002-211557, Apr. 2002. 

24. Von Glahn, U.H., and Gray, V.H., “Effect of Ice Formations on Section Drag of Swept NACA 63A009 Airfoil With Partial-Span Leading-Edge Slat for Various Modes of Thermal Ice Protection,” 

NACA RM E53J30, Mar. 1954. 

25. Gray, V.H., and Von Glahn, U.H., “Heat Requirements for Ice Protection of a Cyclically Gas-Heated, 

36° Swept Airfoil with Partial Span Leading-Edge Slat,” NACA RM E56B23, May 1956. 

26. Koven, W., and Graham, R.R., “Wind-Tunnel Investigation of High-Lift and Stall-Control Devices 

on a 37 Sweptback Wing of Aspect Ratio 6 at High-Reynolds Numbers,” NACA RM L8D29, Sept. 

1948. 

27. Van Hengst, J., Gent, R., Hammond, D., Seubert, R., and Wagner, B., “Ice Accretion and Its Effects 

on Aircraft,” Paper 3 in “Ice Accretion Simulation” AGARD-AR-344, Dec. 1997. 

28. Brumby, R.E., “The Effect of Wing Ice Contamination on Essential Flight Characteristics,” Paper 3 in 

“Effects of Adverse Weather on Aerodynamics,” AGARD-CP-496, Apr.-May, 1991. 

29. Kind, R.J., and Lawrysyn, M.A., “Effects of Frost on Wing Aerodynamics and Take-Off 

Performance,” Paper 8 in “Effects of Adverse Weather on Aerodynamics,” AGARD-CP-496, Dec. 

1991. 
NASA/TM—2013-216381 44 

30. Papadakis, M., Yeong, H.W., Chandrasekharan, R., Hinson, M., and Ratvasky, T.P., “Effects of 

Roughness on the Aerodynamic Performance of a Business Jet Tail,” AIAA Paper 2002-0242, Jan. 

2002. 

31. Bragg, M.B., Kerho, M.F., and Khodadoust, A., “LDV Flowfield Measurements on a Straight and 

Swept Wing With a Simulated Ice Accretion,” AIAA Paper 93-0300, Jan. 1993. 

32. Kwon, O.J., and Sankar, L.N., “Numerical Study of the Effects of Icing on Fixed and Rotary Wing 

Performance,” AIAA Paper 91-0662, Jan. 1990. 

33. Khodadoust, A., and Bragg, M.B, “Measured Aerodynamic Performance of a Swept Wing With a 

Simulated Ice Accretion,” AIAA Paper 90-0490, Jan. 1990. 

34. Papadakis, M., Alansatan, S., and Yeong, H.W., “Aerodynamic Performance of a T-tail With 

Simulated Ice Accretions,” AIAA Paper 2000-0363, Jan. 2003. 

35. Wright, W.B., “User’s Manual for LEWICE Version 3.2,” NASA/CR—2008-214255, Jan. 2008. 

36. 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. 

37. Papadakis, M., Yeong, H.W., Wong, S, Vargas, M and Potapczuk, M., “Experimental Investigation of 

Ice Accretion Effects on a Swept Wing,” DOT/FAA/AR-05/39, Aug. 2005. 

38. Bragg, M.B., and Gregorek, G.M., “Wind Tunnel Investigation of Airfoil Performance Degradation 

Due to Icing,” AIAA Paper 82-0582, Jan. 1982. 

39. Lee, S., and Bragg, M.B., “Effects of Simulated Spanwise Ice Shapes on Airfoils: Experimental 

Investigation,” AIAA Paper 99-0092, Jan. 1999. 

40. Lee, S. and Bragg, M.B., “The Effect of Ridge-Ice Location and the Role of Airfoil Geometry,” 

AIAA Paper 2001-2481, Jun. 2001. 

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

Investigation of Spanwise Step Ice on Airfoil Aerodynamics,” AIAA Paper 98-0490, Jan. 1998. 

42. Johnson, C.L., “Wing Loading, Icing and Associated Aspects of Modern Transport Design,” Journal 

of The Aeronautical Sciences, Vol. 8, No. 2, 1940. 

43. Lee, S., Ratvasky, T.P., Dickes, E., and Thacker, M., “Dynamic Wind-Tunnel Testing of a Sub-Scale 

Iced Business Jet,” AIAA Paper 2006-0261, Jan. 2006. 

44. Papadakis, M., Yeong, H.W., and Wong, S.C., “Aerodynamic Performance of a Swept Wing With 

Simulated Ice Shapes,” AIAA Paper 2004-0734, Jan. 2004. 

45. Busch, G.T, “Experimental Study of Full-Scale Iced-Airfoil Aerodynamic Performance Using SubScale Simulations,” Ph.D. Dissertation, Dept. of Aeronautical and Astronautical Engineering, Univ. 

of Illinois Urbana-Champaign, Urbana, IL, 2009. 

46. Whalen, E.A. “Aerodynamics of Runback Ice Accretions,” Ph.D. Dissertation, Dept. of Aeronautical 

and Astronautical Engineering, Univ. of Illinois Urbana-Champaign, Urbana, IL, 2007. 

47. Papadakis, M., and Gile-Laflin, B.E., “Aerodynamic Performance of a Tail Section with Simulated 

Ice Shapes and Roughness,” AIAA Paper 2001-0539, Jan. 2001.