DOT/FAA/AR-05/39 

Office of Aviation Research 

Washington, D.C. 20591

Experimental Investigation of Ice 

Accretion Effects on a Swept Wing

August 2005 

Final Report 

This document is available to the U.S. public 

through the National Technical Information 

Service (NTIS), Springfield, Virginia 22161. 

U.S. Department of Transportation 

Federal Aviation Administration
NOTICE 

This document is disseminated under the sponsorship of the U.S. 

Department of Transportation in the interest of information exchange. The 

United States Government assumes no liability for the contents or use 

thereof. The United States Government does not endorse products or 

manufacturers. Trade or manufacturer's names appear herein solely 

because they are considered essential to the objective of this report. This 

document does not constitute FAA certification policy. Consult your local 

FAA aircraft certification office as to its use. 

This report is available at the Federal Aviation Administration William J. 

Hughes Technical Center's Full-Text Technical Reports page: 

actlibrary.tc.faa.gov in Adobe Acrobat portable document format (PDF). 
 Technical Report Documentation Page

1. Report No.

DOT/FAA/AR-05/39

2. Government Accession No. 3. Recipient's Catalog No.

 4. Title and Subtitle 

EXPERIMENTAL INVESTIGATION OF ICE ACCRETION EFFECTS ON A 

5. Report Date 

August 2005

SWEPT WING 6. Performing Organization Code

7. Author(s)

Michael Papadakis1, Hsiung-Wei Yeong1, See-Cheuk Wong1, Mario Vargas2, and 

Mark Potapczuk2

8. Performing Organization Report No. 

9. Performing Organization Name and Address

1

Department of Aerospace Engineering 

Wichita State University 

Wichita, KS 67260

10. Work Unit No. (TRAIS) 

 

2

NASA John H. Glenn Research Center at Lewis Field 

Cleveland, OH 44135 

11. Contract or Grant No.

 

12. Sponsoring Agency Name and Address

U.S. Department of Transportation 

Federal Aviation Administration

13. Type of Report and Period Covered

 Final Report

Office of Aviation Research 

Washington, DC 20591

14. Sponsoring Agency Code

 AIR-100

15. Supplementary Notes 

The FAA William J. Hughes Technical Center Technical Monitor was James Riley.

16. Abstract 

An experimental investigation was conducted to study the effects of 2-, 5-, 10-, and 22.5-min ice accretions on the aerodynamic

performance of a swept finite wing. The ice shapes tested included castings of ice accretions obtained from icing tests at the

NASA Glenn Icing Research Tunnel (IRT) and simulated ice shapes obtained with the LEWICE 2.0 ice accretion code. The 

conditions used for the icing tests were selected to provide five glaze ice shapes with complete and incomplete scallop features 

and a small rime ice shape. The LEWICE ice shapes were defined for the same conditions as those used in the icing tests. All 

aerodynamic performance tests were conducted in the 7- x 10-ft Low-Speed Wind Tunnel Facility at Wichita State University. 

Six component force and moment measurements, aileron hinge moments, and surface pressures were obtained for a Reynolds 

number of 1.8 million based on mean aerodynamic chord and aileron deflections in the range of -15o to 20o. Tests were performed 

with the clean wing, six IRT ice shape castings, seven smooth LEWICE ice shapes, and seven rough LEWICE ice shapes. 

Roughness for the LEWICE ice shapes was simulated with 36-size grit. The experiments conducted showed that the glaze ice 

castings reduced the maximum lift coefficient of the clean wing by 11.5% to 93.6%, while the 5-min rime ice casting increased 

maximum lift by 3.4%. Minimum iced wing drag was 133% to 3533% greater with respect to the clean case. The drag of the 

iced wing near the clean wing stall angle of attack was 17% to 104% higher than that of the clean case. In general, the aileron 

remained effective in changing the lift of the clean and iced wings for all angles of attack and aileron deflections tested. Aileron 

hinge moments for the iced wing cases remained within the maximum and minimum limits defined by the clean wing hinge 

moments. Tests conducted with the LEWICE ice shapes showed that in general the trends in aerodynamic performance 

degradation of the wing with the simulated ice shapes were similar to those obtained with the IRT ice shape castings. However, in 

most cases, the ice castings resulted in greater aerodynamic performance losses than those obtained with the LEWICE ice shapes. 

For the majority of the LEWICE ice shapes, the addition of 36-size grit roughness to the smooth ice shapes increased 

aerodynamic performance losses. 

17. Key Words 

Ice accretion, Three-dimensional, Finite swept wing, 

Experimental aerodynamic effects, Ice shape castings, glaze 

ice 

18. Distribution Statement 

This document is available to the public through the National 

Technical Information Service (NTIS) Springfield, Virginia 

22161.

19. Security Classif. (of this report) 

 Unclassified

20. Security Classif. (of this page) 

 Unclassified

21. No. of Pages 

 207 

22. Price

Form DOT F1700.7 (8-72) Reproduction of completed page authorized
ACKNOWLEDGEMENTS 

This work was supported by a grant from the Federal Aviation Administration (FAA). NASA 

Glenn Research Center provided extensive support for this program. The authors would like to 

thank Dr. James T. Riley, the grant technical monitor of the FAA William J. Hughes Technical 

Center, Mr. Eugene Hill, the FAA Icing National Resource Specialist and Tom Bond, the Chief 

of the Icing Branch at the NASA Glenn Research Center for their support. The authors 

acknowledge the extensive efforts of the following NASA personnel who helped with the icing 

tunnel tests and the fabrication of the ice shape castings: Dave Sheldon, Gene Addy, David 

Brinkman, the icing research tunnel crew, and Robert Reminder. Thanks are also due to the 

Raytheon Aircraft Company in Wichita, KS, for providing the wing body mount for the 

aerodynamic tests. Finally, the authors would like to acknowledge the extensive efforts of the 

following Wichita State University (WSU) personnel: Dennis Regnier, Norma Campos, Michael 

Boccia, and Joshua Fong for their assistance with the wind tunnel tests and with the analysis of 

the experimental data; Art Porter of the WSU Machine Shop for his efforts in fabricating the 

wind tunnel model and the simulated ice shapes; the wind tunnel crew for their support during 

the experimental investigation. 

iii/iv
TABLE OF CONTENTS 

Page 

EXECUTIVE SUMMARY xxi 

1. INTRODUCTION 1-1 

2. EXPERIMENTAL FACILITIES AND PROCEDURES 2-1 

2.1 Icing Tests at the NASA Glenn IRT 2-1 

2.1.1 Test Facility 2-1 

2.1.2 Test Model 2-2 

2.1.3 Test Model Hardware for the Icing Tests 2-3 

2.1.4 Icing Conditions and Test Matrix 2-5 

2.1.5 Procedure to Obtain Ice Accretions 2-6 

2.1.6 Casting Process 2-8 

2.2 Aerodynamic Tests at the WSU 7- x 10-ft Tunnel 2-11 

2.2.1 Test Facility 2-11 

2.2.2 Test Model and Installation 2-11 

2.2.3 Icing Research Tunnel Ice Castings and LEWICE Ice Shapes 2-14 

2.2.4 Test Procedure 2-27 

2.2.5 Test Measurements 2-31 

3. RESULTS AND DISCUSSION 3-1 

3.1 Icing Conditions and Ice Accretions (NASA IRT) 3-1 

3.1.1 Icing Condition 1—Complete Scallop Condition (τ = 10 min) 3-1 

3.1.2 Icing Condition 2—Incomplete Scallop Condition (τ = 10 min) 3-2 

3.1.3 Icing Condition 3—Scaled Condition (τ = 5 min) 3-2 

3.1.4 Icing Condition 4—Complete Scallop Condition (τ = 2 min) 3-5 

3.1.5 Icing Condition 5—Complete Scallop Condition (τ = 22.5 min) 3-6 

3.1.6 Icing Condition 6—IPS Failure Condition (τ = 22.5 min) 3-7 

3.2 Aerodynamic Performance Studies 3-8 

3.2.1 Data Quality and Repeatability 3-10 

3.2.2 Compressibility Effects 3-17 

3.2.3 Clean and Iced Wing Performance 3-18 

3.2.4 Comparison of IRT and LEWICE Ice Shapes 3-43 

3.2.5 Effect of Aileron Deflection 3-82 

3.2.6 Effect of Large Scallop Features 3-91 

v
3.2.7 Reynolds Number Effects 3-97 

4. CONCLUSIONS 4-1 

4.1 Performance of Clean Wing 4-1 

4.2 Effects of IRT Ice Shape Castings 4-1 

4.3 Definition of LEWICE Ice Shapes 4-2 

4.4 Effects of Smooth LEWICE Ice Shapes 4-2 

4.5 Effects of Rough LEWICE Ice Shapes 4-2 

4.6 Rough Versus Smooth LEWICE Ice Shapes 4-2 

4.7 Icing Research Tunnel Versus Rough LEWICE Ice Shapes 4-3 

4.8 Aileron Performance 4-3

4.9 Reynolds Number Effects 4-3 

5. REFERENCES 5-1 

APPENDICES 

A—Run Log for Icing Tests at NASA Glenn Icing Research Tunnel

B—Run Log for Aerodynamic Investigation at Wichita State University

 

C—Pressure Data From Icing Tests

D—Flow Angularity Studies 

E—Simulated Frost Using Sandpaper

F—Effect of Aileron Deflection (Supplementary)

G—Computation of LEWICE Ice Shape Profiles

H—Comparison of Two- and Three-Dimensional Wing Ice Shape Traces and 

Lift Data

vi
LIST OF FIGURES 

Figure Page 

2-1 NASA Glenn Icing Research Tunnel Planview 2-1 

2-2 Wing Planform 2-2 

2-3 Airfoil Section and Wing Parameters 2-2 

2-4 Wing in the Low Position Showing the RLEs 2-4 

2-5 Specifications of the Top, Middle, and Bottom RLEs 2-4 

2-6 Low, Normal, and High Positions of the Wing in the IRT 2-4 

2-7 Wing in High Position 2-5 

2-8 Wing in Low Position With the Top Leading-Edge Segment Removed 2-5 

2-9 A Heated Copper Plate Was Used to Cut Ice Accretions for Tracing Purposes 2-8 

2-10 Ice Accretion Pencil-Traced on Cardboard 2-8 

2-11 A Sample of a Pencil Tracing of an Ice Accretion 2-8 

2-12 Top RLE Segment With Ice Accretion Inside the Wooden Molding Box 2-9 

2-13 Middle RLE With IRT-CS22 Ice Accretion Attached to the Inside of a Wooden 

Molding Box—Side View 2-9 

2-14 Leading Edge With Ice Accretion Prepared for Making the Mold—Top View 2-9 

2-15 The Mold and Leading Edge Were Placed Inside the Box to Make the Casting 2-10 

2-16a Close-Up of Ice Shape Casting 2-10 

2-16b Close-Up of the Actual Ice Accretion 2-10 

2-17 Wichita State University 7- x10-ft Wind Tunnel Facility 2-11 

2-18 Wing and Streamlined Body in WSU Test Section—Side View 2-12 

2-19 Installation of Clean Wing in WSU Wind Tunnel—Front View 2-12 

2-20 Installation of Clean Wing in WSU Wind Tunnel—Side View 2-12 

2-21 Installation of Clean Wing on Tunnel Turntable—Rear View 2-13 

vii
2-22 Housing for Servomotor, Torque Cell, and LCDT 2-13 

2-23 Close-Up of Wing Root and Dental Dam—Pressure Surface 2-14 

2-24 Close-Up of Wing Root and Dental Dam—Suction Surface 2-14 

2-25 Icing Research Tunnel CS10 Ice Shape Casting Installed on Wing—Side View 2-15 

2-26 Close-Up of IRT-CS10 Ice Shape Casting at Wing Root 2-15 

2-27 Close-Up of Installation of IRT-IS10 Ice Shape Casting on Wing Model 2-16 

2-28 Close-Up of Installation of IRT-CS22 Ice Shape Casting on Wing Model 2-16 

2-29 Installation of IRT-SC5 Ice Shape Casting on Wing Model—Side View 2-16 

2-30 Icing Research Tunnel CS2 Ice Shape Casting on Test Model—Front View 2-16 

2-31 GLC-305 Swept Wing With IRT-CS22 Ice Shape Casting—Front View 2-17 

2-32 Front View of IRT-IPSF22 Ice Shape Casting at Wing Root 2-17 

2-33 Polyurethane Castings of the Six Ice Shapes Obtained From Icing Tests at IRT 2-17 

2-34 Smooth and Rough LEWICE Ice Shapes 2-20 

2-35 Comparison of IRT and LEWICE Ice Shapes for Icing Condition 1 (CS10) 2-21 

2-36 Comparison of IRT and LEWICE Ice Shapes for Icing Condition 2 (IS10) 2-21 

2-37 Comparison of IRT and LEWICE Ice Shapes for Icing Condition 3 (SC5) 2-22 

2-38 Comparison of IRT and LEWICE Ice Shapes for Icing Condition 4 (CS2) 2-22 

2-39 Comparison of IRT and LEWICE Ice Shapes for Icing Condition 5 (CS22) 2-23 

2-40 Comparison of IRT and LEWICE Ice Shapes for Icing Condition 6 (IPSF22) 2-23 

2-41 Details of Ice Shapes Obtained From Icing Condition 2 2-24 

2-42 Comparison of Profile of LR-CS22N and LR-CS22S Ice Shapes—Tip Section 2-24 

2-43 Comparison of Profile of LR-CS22N and LR-CS22S Ice Shapes—Root Section 2-24 

2-44 Installation of LR-CS10 Rough LEWICE Ice Shape on Wing Model 2-25 

2-45 Close-Up of LR-IS10 Ice Shape at Tip of Wing Model 2-25 

2-46 Installation of LR-SC5 Rough LEWICE Ice Shape on Wing Model 2-26 

viii
2-47 Installation of LR-CS2 Rough LEWICE Ice Shape on Wing Model 2-26 

2-48 Installation of LS-CS22N Smooth LEWICE Ice Shape on Wing Model 2-26 

2-49 Installation of LR-CS22N Rough LEWICE Ice Shape on Wing Model 2-26 

2-50 Installation of LS-IPSF22 Ice Shape on Wing Model—Suction Side 2-27 

2-51 Installation of LS-IPSF22 Ice Shape on Wing Model—Pressure Side 2-27 

2-52 Installation of Seven-Hole Flow Probes on Wing Model—Front View 2-29 

2-53 Installation of Seven-Hole Flow Probes on Wing Model—Side View 2-29 

2-54 Close-Up of Seven-Hole Flow Probe Near to Wing Root (16% Semispan) 2-29 

2-55 Close-Up of Seven-Hole Flow Probe Near to Wing Tip (80% Semispan) 2-29 

2-56 Tufts Installation on Clean Wing and Streamlined Body—Pressure Surface 2-30 

2-57 Installation of Yarn Tufts on Iced Wing and Streamlined Body—Suction Surface 2-30 

2-58 Tufts Show That Flow Downstream of Ice Shape Was Separated Near Wing Tip 2-30 

2-59 Tufts Show Complete Flow Separation at Wing Tip 2-30 

3-1 Icing Condition 1, Front View of Ice Accretion on the Top RLE 3-1 

3-2 Icing Condition 1, Side View of Ice Accretion on Pressure Side of the Top RLE 3-1 

3-3 Icing Condition 1, Front View of the Complete Ice Accretion on the Top RLE 3-2 

3-4 Icing Condition 1, Casting From the Ice Accretion on the Top RLE 3-2 

3-5 Icing Condition 2, Front View of Ice Accretion on the Bottom RLE 3-3 

3-6 Icing Condition 2, Side View of Ice Accretion on Suction Side of the Bottom RLE 3-3 

3-7 Icing Condition 2, Front View of the Complete Ice Accretion on the Bottom RLE 3-3 

3-8 Icing Condition 2, Casting From the Ice Accretion on the Bottom RLE 3-3 

3-9 Icing Condition 3, Front View of Ice Accretion on the Bottom RLE 3-4 

3-10 Icing Condition 3, Side View of Ice Accretion on Suction Side of the Bottom RLE 3-4 

3-11 Icing Condition 3, Front View of the Complete Ice Accretion on the Bottom RLE 3-4 

3-12 Icing Condition 3, Casting From the Ice Accretion on the Bottom RLE 3-4 

ix
3-13 Icing Condition 4, Front View of Ice Accretion on the Middle RLE 3-5 

3-14 Icing Condition 4, Side View of Ice Accretion on Suction Side of the Middle RLE 3-5 

3-15 Icing Condition 4, Front View of the Complete Ice Accretion on the Middle RLE 3-5 

3-16 Icing Condition 4, Casting From the Ice Accretion on the Middle RLE 3-5 

3-17 Icing Condition 5, Front View of Ice Accretion on the Middle RLE 3-6 

3-18 Icing Condition 5, Side View of Ice Accretion on Suction Side of the Middle RLE 3-6 

3-19 Icing Condition 5, Front View of the Complete Ice Accretion on the Middle RLE 3-6 

3-20 Icing Condition 5, Casting From the Ice Accretion on the Middle RLE 3-6 

3-21 Icing Condition 6, Front View of Ice Accretion on the Top RLE 3-7 

3-22 Icing Condition 6, Side View of Ice Accretion on Pressure Side of the 

Top RLE 3-7 

3-23 Icing Condition 6, Front View of the Complete Ice Accretion on the 

Top RLE 3-8 

3-24 Icing Condition 6, Casting From the Ice Accretion on the Top RLE 3-8 

3-25 Notation for Pitching Moment, Hinge Moment, and Aileron Deflection Angle 3-8 

3-26 Regions A, B, and C for Hinge-Moment Curve of Clean and Iced Configurations 3-9 

3-27 Repeatability of CL, CD, CM, and CH; Clean Configuration; Re = 1.8x106; 

δA = 0° 3-11 

3-28 Repeatability of Pressure Distributions at 15% Semispan; Clean Configuration; 

Re = 1.8x106

; δA = 0° 3-12 

3-29 Repeatability of Pressure Distributions at 50% Semispan; Clean Configuration; 

Re = 1.8x106

; δA = 0° 3-13 

3-30 Repeatability of Pressure Distributions at 85% Semispan; Clean Configuration; 

Re = 1.8x106

; δA = 0° 3-14 

3-31 Setup to Check External Balance 3-15 

3-32 Setup to Check Hinge-Moment Load Cell 3-15 

3-33 Data Sampling Studies of CL, CD, CM, and CH; Clean Configuration; 

Re = 1.8x106

; δA = 0° 3-16 

x
3-34 Effect of IRT Ice Shapes on CL, CD, CM, and CH; IRT-CS10, IRT-IS10, 

IRT-SC5, IRT-CS2, IRT-CS22, and IRT-IPSF22 Configurations; 

Re = 1.8x106; δA = 0° 3-19 

3-35 Effect of IRT Ice Shapes on Pressure Distributions at 15% Semispan; 

IRT-CS10, IRT-IS10, IRT-SC5, IRT-CS2, IRT-CS22, and IRT-IPSF22 

Configurations; Re = 1.8x106

; δA = 0° 3-20 

3-36 Effect of IRT Ice Shapes on Pressure Distributions at 50% Semispan; 

IRT-CS10, IRT-IS10, IRT-SC5, IRT-CS2, IRT-CS22, and IRT-IPSF22 

Configurations; Re = 1.8x106; δA = 0° 3-21 

3-37 Effect of IRT Ice Shapes on Pressure Distributions at 85% Semispan; 

IRT-CS10, IRT-IS10, IRT-SC5, IRT-CS2, IRT-CS22, and IRT-IPSF22 

Configurations; Re = 1.8x106; δA = 0° 3-22 

3-38 Effect of Smooth LEWICE Ice Shapes on CL, CD, CM, and CH; 

LS-CS10, LS-IS10, LS-SC5, LS-CS2, LS-CS22N, LS-CS22S, and 

LS-IPSF22 Configurations; Re = 1.8x106

; δA = 0° 3-23 

3-39 Effect of Smooth LEWICE Ice Shapes on Pressure Distributions at 15% 

Semispan; LS-CS10, LS-IS10, LS-SC5, LS-CS2, LS-CS22N, LS-CS22S, 

and LS-IPSF22 Configurations; Re = 1.8x106; δA = 0° 3-24 

3-40 Effect of Smooth LEWICE Ice Shapes on Pressure Distributions at 50% 

Semispan; LS-CS10, LS-IS10, LS-SC5, LS-CS2, LS-CS22N, LS-CS22S, 

and LS-IPSF22 Configurations; Re = 1.8x106; δA = 0° 3-25 

3-41 Effect of Smooth LEWICE Ice Shapes on Pressure Distributions at 85% 

Semispan; LS-CS10, LS-IS10, LS-SC5, LS-CS2, LS-CS22N, LS-CS22S, 

and LS-IPSF22 Configurations; Re = 1.8x106

; δA = 0° 3-26 

3-42 Effect of Rough LEWICE Ice Shapes on CL, CD, CM, and CH; LR-CS10, 

LR-IS10, LR-SC5, LR-CS2, LR-CS22N, LR-CS22S, and LR-IPSF22 

Configurations; Re = 1.8x106

; δA = 0° 3-27 

3-43 Effect of Rough LEWICE Ice Shapes on Pressure Distributions at 15% 

Semispan; LR-CS10, LR-IS10, LR-SC5, LR-CS2, LR-CS22N, LR-CS22S, 

and LR-IPSF22 Configurations; Re = 1.8x106; δA = 0° 3-28 

3-44 Effect of Rough LEWICE Ice Shapes on Pressure Distributions at 50% 

Semispan; LR-CS10, LR-IS10, LR-SC5, LR-CS2, LR-CS22N, LR-CS22S, 

and LR-IPSF22 Configurations; Re = 1.8x106

; δA = 0° 3-29 

xi
3-45 Effect of Rough LEWICE Ice Shapes on Pressure Distributions at 85% 

Semispan; LR-CS10, LR-IS10, LR-SC5, LR-CS2, LR-CS22N, LR-CS22S, 

and LR-IPSF22 Configurations; Re = 1.8x106; δA = 0° 3-30 

3-46 Effect of IRT-CS10, LS-CS10, and LR-CS10 Ice Shapes (Icing Condition 1) 

on CL, CD, CM, and CH; Re = 1.8x106

; δA = 0° 3-44 

3-47 Effect of IRT-IS10, LS-IS10, and LR-IS10 Ice shapes (Icing Condition 2) 

on CL, CD, CM, and CH; Re = 1.8x106; δA = 0° 3-45 

3-48 Effect of IRT-SC5, LS-SC5, and LR-SC5 Ice Shapes (Icing Condition 3) 

on CL, CD, CM, and CH; Re = 1.8x106; δA = 0° 3-46 

3-49 Effect of IRT-CS2, LS-CS2, and LR-CS2 Ice Shapes (Icing Condition 4) 

on CL, CD, CM, and CH; Re = 1.8x106

; δA = 0° 3-47 

3-50 Effect of IRT-CS22, LS-CS22N, LS-CS22S, LR-CS22N, and LR-CS22S 

Ice Shapes (Icing Condition 5) on CL, CD, CM, and CH; Re = 1.8x106; δA = 0° 3-48 

3-51 Effect of IRT-IPSF22, LS-IPSF22, and LR-IPSF22 Ice Shapes 

(Icing Condition 6) on CL, CD, CM, and CH; Re = 1.8x106; δA = 0° 3-49 

3-52 Effect of IRT-CS10, LS-CS10, and LR-CS10 Ice Shapes (Icing Condition 1) 

on Pressure Distributions at 15% Semispan; Re = 1.8x106

; δA = 0° 3-50 

3-53 Effect of IRT-CS10, LS-CS10, and LR-CS10 Ice Shapes (Icing Condition 1) 

on Pressure Distributions at 50% Semispan; Re = 1.8x106; δA = 0° 3-51 

3-54 Effect of IRT-CS10, LS-CS10, and LR-CS10 Ice Shapes (Icing Condition 1) 

on Pressure Distributions at 85% Semispan; Re = 1.8x106; δA = 0° 3-52 

3-55 Effect of IRT-IS10, LS-IS10, and LR-IS10 Ice Shapes (Icing Condition 2) 

on Pressure distributions at 15% Semispan; Re = 1.8x106

; δA = 0° 3-53 

3-56 Effect of IRT-IS10, LS-IS10, and LR-IS10 Ice Shapes (Icing Condition 2) 

on Pressure Distributions at 50% Semispan; Re = 1.8x106

; δA = 0° 3-54 

3-57 Effect of IRT-IS10, LS-IS10, and LR-IS10 Ice Shapes (Icing Condition 2) 

on Pressure Distributions at 85% Semispan; Re = 1.8x106

; δA = 0° 3-55 

3-58 Effect of IRT-SC5, LS-SC5, and LR-SC5 Ice Shapes (Icing Condition 3) 

on Pressure Distributions at 15% Semispan; Re = 1.8x106; δA = 0° 3-56 

3-59 Effect of IRT-SC5, LS-SC5, and LR-SC5 Ice Shapes (Icing Condition 3) 

on Pressure Distributions at 50% Semispan; Re = 1.8x106

; δA = 0° 3-57 

xii
3-60 Effect of IRT-SC5, LS-SC5, and LR-SC5 Ice Shapes (Icing Condition 3) 

on Pressure Distributions at 85% Semispan; Re = 1.8x106

; δA = 0° 3-58 

3-61 Effect of IRT-CS2, LS-CS2, and LR-CS2 Ice Shapes (Icing Condition 4) 

on Pressure Distributions at 15% Semispan; Re = 1.8x106; δA = 0° 3-59 

3-62 Effect of IRT-CS2, LS-CS2, and LR-CS2 Ice Shapes (Icing Condition 4) 

on Pressure Distributions at 50% Semispan; Re = 1.8x106

; δA = 0° 3-60 

3-63 Effect of IRT-CS2, LS-CS2, and LR-CS2 Ice Shapes (Icing Condition 4) 

on Pressure Distributions at 85% Semispan; Re = 1.8x106

; δA = 0° 3-61 

3-64 Effect of IRT-CS22, LS-CS22N, LS-CS22S, LR-CS22N, and LR-CS22S 

Ice Shapes (Icing Condition 5) on Pressure Distributions at 15% Semispan; 

Re = 1.8x106

; δA = 0° 3-62 

3-65 Effect of IRT-CS22, LS-CS22N, LS-CS22S, LR-CS22N, and LR-CS22S 

Ice Shapes (Icing Condition 5) on Pressure Distributions at 50% Semispan; 

Re = 1.8x106; δA = 0° 3-63 

3-66 Effect of IRT-CS22, LS-CS22N, LS-CS22S, LR-CS22N, and LR-CS22S 

Ice Shapes (Icing Condition 5) on Pressure Distributions at 85% Semispan; 

Re = 1.8x106; δA = 0° 3-64 

3-67 Effect of IRT-IPSF22, LS-IPSF22, and LR-IPSF22 Ice Shapes (Icing Condition 6) 

on Pressure Distributions at 15% Semispan; Re = 1.8x106

; δA = 0° 3-65 

3-68 Effect of IRT-IPSF22, LS-IPSF22, and LR-IPSF22 Ice Shapes (Icing Condition 6) 

on Pressure Distributions at 50% Semispan; Re = 1.8x106; δA = 0° 3-66 

3-69 Effect of IRT-IPSF22, LS-IPSF22, and LR-IPSF22 Ice Shapes (Icing Condition 6) 

on Pressure Distributions at 85% Semispan; Re = 1.8x106

; δA = 0° 3-67 

3-70 Effect of Aileron Deflection on Lift and Aileron Hinge-Moment Coefficients; 

Clean Configuration; Re = 1.8x106

 3-83 

3-71 Effect of Aileron Deflection on Lift and Aileron Hinge-Moment Coefficients; 

IRT-CS22 Ice Shape; Re = 1.8x106

 3-84 

3-72 Effect of Aileron Deflection on Lift and Aileron Hinge-Moment Coefficients; 

IRT-IPSF22 Ice Shape; Re = 1.8x106

 3-85 

3-73 Effect of Aileron Deflection on Lift and Aileron Hinge-Moment Coefficients; 

LR-IPSF22 Ice Shape; Re=1.8x106

 3-86 

3-74 Ice Accretion on a Swept Wing at Glaze Ice Conditions 3-93 

xiii
3-75 Development of Roughness Elements and Ice Feathers During Ice Accretion 

on a Swept Wing 3-93 

3-76 Icing Research Tunnel CS22 Ice Shape Filled With Modeling Compound 3-94 

3-77 Effect of Scallop Features on Aerodynamic Performance; IRT-CS22 Ice Shape 

With and Without Modeling Compound; Re = 1.8x106

; δA = 0° 3-95 

3-78 Reynolds Number Effect on CL; Clean and Iced Wing; δA = 0° 3-98 

xiv
LIST OF TABLES 

Table Page 

2-1 Icing Conditions for IRT Ice Shape Castings 2-6 

2-2 Summary of Ice Shape Sections and Corresponding Icing Conditions 2-7 

2-3 Icing Conditions for LEWICE Ice Shapes 2-15 

2-4 Test Conditions for Aerodynamic Performance Data 2-28 

3-1 Effect of IRT Ice Shapes on CL and CD; ReMAC = 1.8x106; δA = 0° 3-31 

3-2 Effect of IRT Ice Shapes on CL,stall, αstall, and CDmin; ReMAC = 1.8x106; δA = 0° 3-31 

3-3 Effect of IRT Ice Shapes on Hinge-Moment Coefficient Slope; 

ReMAC = 1.8x106; δA = 0° 3-32 

3-4 Effect of Smooth LEWICE Ice Shapes on CL and CD; ReMAC = 1.8x106; 

δA = 0° 3-32 

3-5 Effect of Smooth LEWICE Ice Shapes on CL,stall, αstall, and CD,min; 

ReMAC = 1.8x106; δA = 0° 3-33 

3-6 Effect of Smooth LEWICE Ice Shapes on Hinge-Moment Coefficient Slope; 

ReMAC = 1.8x106, δA = 0° 3-33 

3-7 Effect of Rough LEWICE Ice Shapes on CL and CD; ReMAC = 1.8x106; 

δA = 0° 3-34 

3-8 Effect of Rough LEWICE Ice Shapes on CL,stall, αstall, and CD,min; 

ReMAC = 1.8x106; δA = 0° 3-34 

3-9 Effect of Rough LEWICE Ice Shapes on Hinge-Moment Coefficient Slope; 

ReMAC = 1.8x106

; δA = 0° 3-35 

3-10 Cl,stall of all Configurations With Aileron Deflection; ReMAC = 1.8x106 3-87 

xv
LIST OF ACRONYMS AND SYMBOLS 

2D Two-dimensional 

3D Three-dimensional 

Al Aluminum 

AOA Angle of attack (in degrees) 

CFD Computational fluid dynamics 

FAA Federal Aviation Administration 

FAR Federal Aviation Regulation 

IPS Ice protection system 

IPSF Ice protection system failure condition, 22.5-min ice accretion time (icing 

condition 6) 

IRT Icing Research Tunnel 

LE Leading edge 

LEB Leading-edge bubble 

LEWICE NASA Glenn Ice Accretion Code (Version 2.0) 

LS Wing lower surface 

LTPT Low-Turbulence Pressure Tunnel 

LVDT Linear variable differential transducer 

LWC Liquid water content 

MAC Mean aerodynamic chord 

MVD Median volumetric diameter 

NASA National Aeronautics and Space Administration 

PSI Pressure Systems Incorporated 

xvi
RLE Removable leading edge 

TE Trailing edge 

US Wing upper surface 

WSU Wichita State University 

a0 2D lift curve slope (per radian) 

AR Aspect ratio 

b Wing span 

Ca Aileron mean chord length behind the hinge line 

CD Drag coefficient 

CD,min Minimum drag coefficient 

ΔCD Percentage change in drag coefficient with respect to the clean wing 

CH Hinge-moment coefficient 

CH,α = dCH/dα; Change in hinge-moment coefficient with respect to AOA; control 

floating parameter 

CH,δ = dCH/dδ; slope of CH vs. δA curve; control heaviness parameter 

CL Lift coefficient 

CL,max Maximum lift coefficient 

CL,stall Lift coefficient corresponding to the point on the lift curve where massive 

flow separation (stall) occurs 

Δ CL Percentage change in lift coefficient with respect to the clean wing 

CL,α = dCL/dα; change in lift coefficient with respect to AOA 

CM Pitching-moment coefficient 

CM,clean Pitching-moment coefficient of the clean wing about the 25% MAC location 

xvii
CM,LEWICE Pitching-moment coefficient of wing with LEWICE ice shape about the 25% 

MAC location 

CM,MAC/4 Pitching-moment coefficient about the 25% MAC location 

dCM/dα Change in pitching-moment coefficient with respect to the AOA 

Cp Pressure coefficient 

Cp0 Suction pressure coefficient for incompressible flow 

Cpcr Critical pressure coefficient 

dcr Critical distance 

HM Hinge moment 

M∞ Free-stream Mach number 

S Wing planform area 

Sa Aileron planform area behind the hinge line 

ReMAC Reynolds number based on MAC 

Ttotal Total temperature 

V∞ Free-stream velocity 

x/c Nondimensional chordwise distance 

α Geometric angle of attack (in degrees) 

αe Effective angle of attack (in degrees) 

αlocal Local geometric angle of attack (in degrees) 

αstall Geometric angle of attack at which stall occurs (in degrees) 

δA Aileron deflection angle (in degrees) 

γ = 1.4; ratio of specific heats 

xviii
Λ Sweepback angle at leading edge 

ρ Free-stream density 

τ Ice accretion time (in minutes) 

xix/xx
EXECUTIVE SUMMARY

Aircraft certification requires the evaluation of the effects of ice accretions on aircraft 

aerodynamic components such as wings, tails, and control surfaces. Ice accumulation on 

aerodynamic surfaces can have a significant impact on aircraft performance, handling qualities, 

and thus, aircraft safety. In general, ice accumulation on aerodynamic surfaces can cause flow 

separation whose extent over the aerodynamic surface is a function of ice shape and wing 

geometry. The term critical is often used to identify ice shapes responsible for large degradation 

in the aerodynamic performance of lifting surfaces. 

A number of experimental studies have been conducted over the years in an effort to assess the 

effect of various forms of ice accretions on aircraft aerodynamic performance and handling 

qualities. Most of these studies, however, have been limited to two-dimensional airfoil sections 

due to tunnel and model cost constraints. Only a small number of investigations have addressed 

the impact of ice shapes on iced three-dimensional (3D) finite wings. 

To address the lack of experimental data for iced 3D finite wing configurations, the Federal 

Aviation Administration (FAA), Wichita State University (WSU), the National Aeronautics and 

Space Administration (NASA), and general aviation aircraft manufacturers initiated a 

collaborative research program in the fall of 2000 for a systematic evaluation of ice accretion 

effects on finite wings. The main objective of this research program was the development of a 

3D experimental database of ice accretion effects on a swept 3D finite wing. The database can 

be used for developing certification guidance material and for improving and validating 

simulation tools for aerodynamic analysis and design. A research grant was awarded to WSU to 

design and fabricate a wind tunnel model and to conduct icing and aerodynamic performance 

tests to develop the required database. 

The specific goals of the research program were to (1) investigate wing sensitivity to various 

forms of ice accretions, (2) compare the effects of actual ice shapes from icing tunnel tests with 

equivalent simulated ice shapes, (3) assess the effects of glaze ice shape features such as horn 

angle, horn height and surface roughness on wing aerodynamic performance, and (4) develop an 

experimental database of ice shape effects on the aerodynamic performance of a swept finite 

span wing with an aileron control surface. The principal accomplishments of this research 

program, which was completed in the fall of 2002, are summarized below. 

A number of meetings were conducted with the FAA, NASA, and general aviation 

manufacturers to select a wing model, ice shapes, and test conditions, and to define experimental 

methodologies and procedures. A 5-ft semispan swept finite wing reflection plane model 

equipped with an aileron control surface was designed, fabricated, and instrumented at WSU. 

Tests were conducted at the NASA Glenn Icing Research Tunnel (IRT) to generate five glaze ice 

shapes with complete and incomplete scallop features and one rime ice shape. Ice shape castings 

were produced by NASA personnel from the actual ice accretions for aerodynamic testing. 

Simulated 3D ice shapes were defined with the NASA Glenn LEWICE ice accretion code for the 

same icing conditions used in the IRT icing tests. The 3D ice shapes were fabricated out of 

wood or aluminum at WSU and were prepared for aerodynamic testing. Extensive experiments 

were conducted at the WSU 7- x 10-ft wind tunnel facility over a period of 10 weeks to generate 

xxi
the required data for the clean and iced wing. Twenty ice shapes and eight roughness cases were 

tested along with the clean wing. The ice shapes included six IRT castings and seven smooth 

and seven rough LEWICE shapes. The roughness cases included 120- and 150-grit sandpaper to 

simulate the effect of frost on wing aerodynamic performance. Lift, drag, pitching moment, 

hinge moment, and pressure distributions were obtained for all configurations tested. 

The experimental results obtained showed that the stall lift coefficients for the wing with the 

glaze ice shape castings were 11.5% to 93.6% less than the clean wing. For the 5-min rime ice 

shape, the stall lift coefficient was 3.4% higher than the clean wing. The IRT ice shape castings 

tested increased the minimum clean wing drag coefficient by 133% to 3533% and increased drag 

near stall by 17% to 104%. In general, the aileron remained effective in changing the lift of the 

clean and iced wings for all angles of attack and aileron deflections tested. Aileron hinge 

moments for the iced wing remained within the maximum and minimum limits defined by the 

clean wing hinge-moment data. In general, the trends in aerodynamic performance degradation 

of the wing with the simulated rough LEWICE ice shapes were similar to those obtained with the 

IRT ice shape castings. However, in most cases, the ice shape castings resulted in greater 

aerodynamic performance losses than that obtained with the rough LEWICE shapes. In most 

cases, the rough LEWICE ice shapes caused greater aerodynamic performance degradation than 

did their smooth counterparts. 

xxii
1. INTRODUCTION. 

Aircraft certification requires the evaluation of the effects of ice accretions on aircraft 

aerodynamic components such as wings, tails, and control surfaces. Ice accumulation on 

aerodynamic surfaces can have a significant impact on aircraft performance, handling qualities, 

and thus, aircraft safety. A wide range of ice accretions is possible, depending on aircraft 

configuration, icing, and flow conditions. Potential ice accretions include glaze ice, rime ice, 

runback, and beak ice, as well as small ice shapes which can have considerable degradation in 

aircraft performance. In general, ice accumulation on aerodynamic surfaces can cause flow 

separation whose extent over the aerodynamic surface is a function of ice shape and wing 

geometry. The term critical is often used to identify ice shapes responsible for large degradation 

in the aerodynamic performance of lifting surfaces. 

A number of experimental studies have been conducted over the years (refer to reference 1 for a 

comprehensive review) in an effort to assess the effect of various forms of ice accretions on 

aircraft aerodynamic performance and handling qualities. Most of these studies have been 

limited to two-dimensional (2D) airfoil sections due to tunnel and model cost constraints. Only a 

small number of investigations have addressed the impact of ice shapes on finite wings. 

To address the lack of experimental data for three-dimensional (3D) iced wing configurations, 

scientists and engineers from the Federal Aviation Administration (FAA), the National 

Aeronautics and Space Administration (NASA), the general aviation aircraft industry, and 

Wichita State University (WSU) conducted a meeting to plan a research program for a 

systematic evaluation of ice accretion effects on finite wings. The main objective of this 

research program was the development of a 3D experimental database of ice accretion effects on 

a swept finite wing. The database was needed for developing certification guidance material and 

for improving and validating simulation tools for aerodynamic analysis and design. A research 

grant was awarded to WSU to design and fabricate a wind tunnel model and to conduct icing and 

aerodynamic performance tests to develop the required database. The specific goals of the WSU 

research effort were to 

• investigate wing sensitivity to various forms of ice accretions. 

• compare the effects of actual ice shapes from icing tunnel tests with equivalent simulated 

ice shapes. 

• assess the effects of glaze ice shape features such as horn angle, horn height, and surface 

roughness on wing aerodynamic performance. 

• develop an experimental database of ice shape effects on aerodynamic performance of a 

swept finite wing with an aileron control surface. 

To accomplish the above objectives, wind tunnel tests were planned with a wing representative 

of modern business jet and regional jet aircraft wing planforms. The wing selection was based 

on input from the FAA, NASA, and the aircraft industry. An important consideration in the 

selection of the airfoil section for the finite wing was the availability of 2D iced airfoil 

1-1
aerodynamic performance data for comparison with the 3D iced wing aerodynamic performance 

results. The airfoil selected was a GLC-305 section, which is representative of a modern 

business jet wing section. NASA has conducted extensive 2D wind tunnel tests with this airfoil 

in recent years using a range of ice accretions. 

The ice shapes selected for the experimental investigation included the following: 

• Polyurethane castings of ice accretions obtained from icing tests at the NASA Glenn 

Icing Research Tunnel (IRT). 

• Smooth LEWICE ice shapes obtained for the same icing conditions as the ones used in 

the generation of the ice shape castings. 

• Rough LEWICE ice shapes obtained by adding grit roughness to the surface of the 

smooth LEWICE ice shapes. 

This report describes the icing tests performed in the IRT and the aerodynamic tests 

conducted at WSU with the ice castings and the simulated LEWICE ice shapes. 

Experimental data presented includes lift, drag, pitching moment, aileron hinge moment, 

and surface pressure data. 

1-2
2. EXPERIMENTAL FACILITIES AND PROCEDURES. 

The experimental investigation consisted of ice accretion tests and aerodynamic performance 

wind tunnel tests. The icing tests were conducted over a time period of 5 days at the IRT facility 

to obtain six ice shape castings for aerodynamic testing. The aerodynamic performance 

investigation took place at the 7- x 10-ft Low-Speed Wind Tunnel facility at WSU. The clean 

wing and 20 ice shape configurations were tested at the WSU wind tunnel over a time period of 

10 weeks. This section describes the test model and ice shapes tested, the experimental setup, 

test conditions, test measurements, and procedures used in the experimental investigations 

conducted. 

2.1 ICING TESTS AT THE NASA GLENN IRT. 

The first part of the experiment was conducted at the NASA Glenn 6- x 9-ft Icing Research 

Tunnel to obtain castings of ice accretions formed on a swept wing model. Details of the icing 

tests are provided in the following sections. 

2.1.1 Test Facility. 

The IRT is a closed-loop refrigerated wind tunnel. Its test section is 6 ft (1.8 m) high, 9 ft 

(2.7 m) wide, and 20 ft (6.0 m) long. In the test section, the total air temperature can be varied 

between -20°F (-30°C) and +33°F (+1°C), within accuracy of ±1°F (±0.5°C), and a maximum 

velocity of 390 mph (160 m/s) can be attained. A spray system allows control of the liquid water 

content (LWC) between 0.2 to 3.0 g/m3 and provides droplet median volumetric diameters 

(MVD) from 15 to 40 μm. Figure 2-1 shows the planview of the NASA Glenn Icing Research 

Tunnel. 

FIGURE 2-1. NASA GLENN ICING RESEARCH TUNNEL PLANVIEW 

2-1
2.1.2 Test Model. 

Aerodynamic considerations and facility size limitations determined the overall size of the wing 

model. Details of the swept wing model are provided in figures 2-2 and 2-3. The model was a 

swept finite wing with a GLC-305 airfoil section aligned in the streamwise direction. The airfoil 

section had a maximum thickness to chord ratio of 0.087 at approximately 38% local chord. The 

airfoil section remained constant from the wing root to the wing tip. The wing had a 28° leading 

edge (LE) sweep, a 15.6° trailing-edge (TE) sweep, a 60-in. semispan, a 7.35 ft2 area (half wing), 

an aspect ratio (AR) of 6.80 (left and right wing), a taper ratio of 0.4, and a geometric twist of 0°

at the root and -4° (washout) at the tip. The wing root and tip chords were 25.2 inches and 10.08 

inches respectively. In addition, the wing mean aerodynamic chord (MAC) was 18.72 inches 

and was located 25.74 inches from the wing root. The model was instrumented with 203 

pressure ports distributed chordwise at five spanwise locations corresponding to 15%, 30%, 

50%, 68%, and 85% semispan. 

AIRFOIL SECTION 

1. GLC-305 section in streamwise direction (t/cmax=0.087 at 

x/c=0.38). Airfoil section is constant from root to tip. 

2. Trailing edge thickness = 0.04 in (constant spanwise) 

 

WING PLANFORM 

1. 5 ft semispan (72% of WSU tunnel height) 

2. Root chord = 25.2 in, Tip chord = 10.08 in, MAC = 18.72 in 

3. t/c varies linearly from root to tip 

4. Planform Area (half wing, S/2) = 7.35 ft2

; wing area to test 

section area ratio = 7.35/(7x10) = 0.105 

5. taper ratio = 0.4, trapezoidal planform, linear chord variation 

from root to tip 

6. AR = b2

/S = 102 /(2*7.35) = 6.8027 

7. 25° quarter chord sweep, 28° leading edge sweep 

WING GEOMETRIC TWIST AND DIHEDRAL 

1. Geometric twist about 25% local chord (0° at root, -4° at tip) 

2. Dihedral = 0°AILERON CONTROL SURFACE 

1. 25% local chord (from 70% to 95% semispan); 

Croot = 3.66 in, Ctip = 2.712 in

2. Elliptical leading edge (pivot at 80% local wing chord), 

Gap = 0.5% local chord

3. Deflection range: -20° to +20°

4. Aileron area behind the hinge line, Sa = 0.2647 ft2

0.00 0.10 0.20 0.30 0.40 0.50 0.60 0.70 0.80 0.90 1.00

-0.10

-0.05

0.00

0.05

0.10

25.200 

280

60.000

15.000 

42.000 

 10.080 

2.920 

2.163 

All dimensions are in inches 

 MAC 

 25% chord line 

18.720 

25.740 

5. Aileron average chord behind the hinge line, 

ca = [(2.163+2.92)/2/12] = 0.2118 ft

FIGURE 2-2. WING PLANFORM FIGURE 2-3. AIRFOIL SECTION AND WING 

PARAMETERS 

2-2
2.1.3 Test Model Hardware for the Icing Tests. 

The icing tests conducted for the purpose of generating ice castings for the swept finite wing 

required considerable planning to maximize the number of test runs during each day of testing. 

In general, once an icing test was completed, the ice casting process required approximately 1 

day to provide a mold. Thus, the wing cannot be used until the mold was completed and 

removed from the wing. Another limitation of the ice casting process was the spanwise length of 

the casting cannot exceed 25 inches. 

The goal of the icing tests conducted in the IRT facility with the 60-in. semispan finite wing was 

to produce six ice castings approximately 68 inches long (the length of wing measured along the 

swept LE) in 5 days of testing. Due to the limitations of the ice casting process, only five 25-in. 

segments could be produced in 5 days, unless significant modifications were made to the wing 

model to allow multiple ice castings during each test day. This was accomplished by designing 

and constructing four wing leading edges. The first leading edge was for aerodynamic 

measurements only and extended the full span of the wing. Leading edges two through four, 

however, were divided into three spanwise segments each, as shown in figures 2-4 and 2-5. 

Thus, a total of nine leading-edge segments were produced in addition to the full-span leading 

edge. The removable leading edges were labeled top, middle, and bottom. The bottom 

removable leading edge (RLE) was 25 inches long and extended from the wing root to 36.79% 

semispan. The middle segment was also 25 inches long and extended from 36.79% semispan to 

73.78% semispan. The top RLE had a length of 17.95 inches and extended from 73.78% 

semispan to the wing tip. Note that the sum of the lengths of the three RLEs was 67.98 inches, 

which is the distance from the wing root to the wing tip measured along the swept leading edge. 

This distance is equal to the wing semispan divided by the cosine of the leading-edge sweep 

angle (i.e., 28 cos °

60 ). The advantage of having 18 leading-edge segments was that every time an 

ice accretion test was completed, the leading-edge segment with the ice shape was removed for 

the ice molding process and a clean segment was placed on the wing. This permitted multiple 

icing runs to be conducted each test day. 

Additional model hardware had to be designed and fabricated to support the icing tests at the 

NASA Glenn IRT. The additional hardware included brackets and a 1-ft wing extension to 

permit the placement in the wing from 5.5 inches below the IRT floor to 11.5 inches above the 

floor. This was required to place the wing inside the uniform LWC region of the icing cloud. 

By lowering the wing below the tunnel floor, LWC uniformity was maintained over the middle 

and top segments of the wing. With the wing in the high position, the middle and bottom 

segments of the wing leading edge were exposed to the uniform region of the icing cloud. For 

aerodynamic pressure measurement, it was necessary to install the wing with its root section 

placed on the tunnel floor. This required additional hardware to be fabricated. The three wing 

placements with respect to the IRT tunnel floor are shown in figures 2-6 to 2-8. 

2-3
All dimensions are in inches 

 MAC 44.15

22.07

19.30 

13.84 

25.00

25.00

17.95 

Bottom RLE

Middle RLE

Top RLE 

Cut C

Cut B 

Cut A 

Top Removable 

Leading Edge

Middle 

Removable 

Leading Edge

Bottom Removable 

Leading Edge 

FIGURE 2-4. WING IN THE LOW 

POSITION SHOWING THE RLEs 

FIGURE 2-5. SPECIFICATIONS OF THE 

TOP, MIDDLE, AND BOTTOM RLEs 

a. Low wing position Uniform Cloud Extent 

Tunnel Ceiling

Tunnel Floor

b. Normal wing position c. High wing position 

FIGURE 2-6. LOW, NORMAL, AND HIGH POSITIONS OF THE WING IN THE IRT 

(Cuts were made to remove ice accretion from wing at locations A, B, and C.) 

2-4
Top RLE Removed 

FIGURE 2-7. WING IN HIGH POSITION FIGURE 2-8. WING IN LOW POSITION 

WITH THE TOP LEADING-EDGE 

SEGMENT REMOVED 

2.1.4 Icing Conditions and Test Matrix. 

Table 2-1 lists the six icing conditions selected to obtain the ice accretions for the aerodynamic 

investigation. These icing conditions were selected to provide four glaze ice shapes with 

complete and incomplete scallop features, a 22.5-min glaze ice accretion representative of an ice 

protection system (IPS) failure case, and a 5-min rime ice shape. Glaze ice accretions with 

complete and incomplete scallop features are discussed in detail in references 2 to 4. The four 

ice shapes with scallop features were obtained using progressively longer ice accretion times to 

provide a range of horn sizes for aerodynamic testing. The rime ice shape was based on a scaled 

icing condition from icing tests conducted with a 2D 36-in. chord GLC-305 airfoil. Ice shape 

details are provided in table 2-2. Appendix A lists all of the runs done in the IRT to obtain ice 

accretions for the top, middle, and bottom RLEs at each of the six icing conditions. 

2-5
TABLE 2-1. ICING CONDITIONS FOR IRT ICE SHAPE CASTINGS 

Icing 

Condition Description 

*AOA 

(deg) 

V 

(mph)

Ttotal

(°F) 

LWC 

(g/m3

) 

MVD 

(μm) 

τ

(min)

1 (Glaze) Complete Scallop Condition 

(ID: Ice1 or IRT-CS10) 4 250.0 25.0 0.68 20.0 10.0 

2 (Glaze) Incomplete Scallop Condition 

(ID: Ice2 or IRT-IS10) 4 150.0 25.0 0.65 20.0 10.0 

3 (Rime) Scaled Condition from 2D tests 

(ID: Ice3 or IRT-SC5) 6 201.3 11.7 0.51 14.5 5.0 

4 (Glaze) Complete Scallop Condition 

(ID: Ice4 or IRT-CS2) 4 250.0 25.0 0.68 20.0 2.0 

5 (Glaze) Complete Scallop Condition 

(ID: Ice5 or IRT-CS22) 4 250.0 25.0 0.68 20.0 22.5 

6 (Glaze) Failed Ice Protection 

(ID: Ice6 or IRT-IPSF22) 4 150.0 27.0 0.46 20.0 22.5 

* AOA = angle of attack 

2.1.5 Procedure to Obtain Ice Accretions. 

Two icing runs at each icing condition were needed to obtain one complete ice accretion over the 

total length of the wing leading edge (i.e., 67.95 inches). The first run with the wing in the low 

position provided the ice accretion on the top RLE. In this position, the ice accretions on the 

middle and bottom RLEs were not used. A second run with the wing in the high position 

provided the ice accretion on the middle and bottom RLEs. In this position, the ice accretion of 

the leading edge near the tip was not used. From the two wing positions (at a given icing 

condition), three removable leading edges with ice accretions were obtained and were used to 

make a set of three molds, and later a set of three castings. Each icing condition resulted in a set 

of three castings, providing a full-span ice shape casting. Ultimately, six sets of castings were 

obtained during the test (one set for each of the six icing conditions) for a total of 18 castings. 

At the start of each run, the tunnel was brought to the target velocity and total temperature before 

activating the tunnel spray system. During the run, the tunnel parameters were recorded using 

the NASA Glenn Escort data acquisition system. Once the target ice accretion time was reached, 

the spray system was shut off, and the tunnel was brought to idle. At this point, tunnel crew 

entered the test section and photographs were taken with a 35-mm camera following a 

predetermined sequence and locations. Cuts were then made on the ice accretion using a heated 

1/8-in.-thick copper plate to prepare the leading edges for tracings and removal, as depicted in 

figure 2-9. When the wing was placed in the low position, only one cut (perpendicular to the 

LE) was needed at the location where the top and the middle leading edges joined. This cut was 

named cut A and is shown in figure 2-5. When the wing was placed in the high position, three 

cuts were made: one at the intersection where the top and middle RLEs joined (cut A), another 

at the location where the middle and bottom RLEs joined (cut B), and one at the lower part of the 

bottom RLE (cut C). Cuts A and B were made perpendicular to the LE, whereas cut C was 

2-6
parallel to the flow direction (streamwise direction) because it was located at the exact root of the 

wing. At each cut, illustrated in figure 2-10, a pencil tracing of the ice shape was made on a 

cardboard template. Figure 2-11 shows an example of a pencil tracing. After the ice section 

tracings were completed, the leading edges were detached from the wing and moved to the cold 

room adjacent to the tunnel to begin the casting process. 

TABLE 2-2. SUMMARY OF ICE SHAPE SECTIONS AND CORRESPONDING 

ICING CONDITIONS 

Ice1 or IRT-CS10 

Complete Scallops 

V = 250 mph 

AOA = 4°

Ttotal = 25 °F 

LWC = 0.68 g/m3

MVD = 20 μm 

Time = 10 min 

Ice2 or IRT-IS10 

Incomplete Scallops 

V = 150 mph 

AOA = 4°

Ttotal = 25 °F 

LWC = 0.65 g/m3

MVD = 20 μm 

Time = 10 min 

Ice3 or IRT-SC5 

No Scallops 

V = 201 mph 

AOA = 6°

Ttotal = 11.7 °F 

LWC = 0.51 g/m3

MVD = 14.5 μm 

Time = 5 min 

Ice4 or IRT-CS2

Initial Formation 

 of Scallops 

V = 250 mph 

AOA = 4°

Ttotal= 25 °F 

LWC = 0.68 g/m3

MVD = 20 μm 

Time = 2 min 

Ice5 or IRT-CS22 

Complete Scallops 

V = 250 mph 

AOA = 4°

Ttotal = 25 °F 

LWC = 0.68 g/m3

MVD = 20 μm 

Time = 22.5 min 

Ice6 or IRT-IPSF22 

IPS Failure Case 

V = 150 mph 

AOA = 4°

Ttotal = 27 °F 

LWC = 0.46 g/m3

MVD = 20 μm 

Time = 22.5 min 

A

A: Normal to LE

B: Normal to LE

C: Streamwise

B

C

A

A: Normal to LE

B: Normal to LE

C: Streamwise

B

C

B

A

C

A

B

C

Ice1 Ice2 Ice3

Ice4 Ice5 Ice6 B

A

C

2-7
FIGURE 2-9. A HEATED COPPER PLATE 

WAS USED TO CUT ICE ACCRETIONS 

FOR TRACING PURPOSES 

Copper Plate 

FIGURE 2-10. ICE ACCRETION PENCILTRACED ON CARDBOARD (IRT-CS22 at 

tip of bottom RLE) 

(IRT-CS22 at tip of bottom RLE) 

FIGURE 2-11. A SAMPLE OF A PENCIL TRACING OF AN ICE ACCRETION 

(IRT-CS22 at root of top RLE) 

2.1.6 Casting Process. 

The casting process involved two steps: making the mold and making the actual casting. 

The process started when the LE was removed and taken to the cold room. The LE was placed 

inside a wooden box (figure 2-12) made specially for the molding process. The top of the box 

and one of its sides were removed to free up space for handling the delicate ice accretion. Once 

the RLE was attached to the inside of the box with bolts, the removed side of the box was 

reinstalled. The box was left open at the top for pouring in the mold material. Figures 2-13 and 

2-8
2-14 show side and top views of the ice accretion on the RLE attached inside the wooden box. 

Styrofoam pieces were placed around the ice accretion, without touching it, to reduce the 

quantity of mold material needed. 

Leading Edge 

with Ice Accretion

FIGURE 2-12. TOP RLE SEGMENT WITH ICE ACCRETION INSIDE THE 

WOODEN MOLDING BOX 

FIGURE 2-13. MIDDLE RLE WITH IRTCS22 ICE ACCRETION ATTACHED TO 

THE INSIDE OF A WOODEN MOLDING 

BOX—SIDE VIEW 

FIGURE 2-14. LEADING EDGE WITH ICE 

ACCRETION PREPARED FOR MAKING 

THE MOLD—TOP VIEW 

2-9
Preparation of the Dow Corning 3110 room temperature vulcanized mold material started before 

the tunnel runs. Dow Corning 3110 was poured in containers and mixed with Dow Corning 200, 

20 centistokes, at 10% of the Dow Corning 3110 by weight. Each container was degassed for 30 

minutes or more in a bell jar and then placed in a freezer. When the leading edge was ready 

inside the box, a container with Dow Corning 3110 was removed from the freezer and Dow 

Corning no. 4 catalyst was added, at a 1-to-125 ratio of the Dow Corning 3110 by weight. After 

mixing, it was degassed in the bell jar for 5 minutes, moved to the cold room, and poured into the 

box containing the leading edge with the ice accretion. The mold was left to cure overnight. In 

the morning, the mold was separated from the leading edge. 

Once the molds were finished, they were moved to the model shop of the NASA Glenn Research 

Center, where the castings were made. To make a casting, a mold was placed inside the same 

box used to make it. The wing RLE that was used to make the mold was placed in the same 

position as when the mold was made. This is illustrated in figure 2-15. A separation distance 

was left between the LE and the mold to create a thin wall for the casting. A polyurethane 

elastomer, Ciba-Geigy RP 6430, was poured in and left to cure. After curing, the LE was 

separated from the casting, and the mold material was removed from around the casting. 

FIGURE 2-15. THE MOLD AND LEADING EDGE WERE PLACED INSIDE THE 

BOX TO MAKE THE CASTING 

Figure 2-16a shows a close-up of the casting obtained from the ice accretion, and figure 2-16b 

shows a close-up of the actual ice accretion. Comparison of the two figures illustrates the detail 

captured with the casting process. 

 

FIGURE 2-16a. CLOSE-UP OF ICE SHAPE 

CASTING 

FIGURE 2-16b. CLOSE-UP OF THE 

ACTUAL ICE ACCRETION 

2-10
2.2 AERODYNAMIC TESTS AT THE WSU 7- x 10-ft TUNNEL. 

All aerodynamic performance tests were conducted in the WSU 7- x 10-ft Beech Memorial LowSpeed Wind Tunnel. Details of the wind tunnel investigation are provided below. 

2.2.1 Test Facility. 

The WSU wind tunnel is a single-return, closed circuit facility with a maximum speed of 

160 mph (235 ft/s), corresponding to a Reynolds number of 1.46 million per foot. The test 

section is 7 ft high by 10 ft wide by 12 ft long. Four screens located in the plenum chamber 

upstream of the test section are used for flow conditioning. The contraction ratio between the 

plenum and test sections is 6 to 1. The tunnel is equipped with a four-bladed, 11-ft diameter, 

variable pitch propeller, which is driven by a 1000 horsepower electric motor. Figure 2-17 

shows the planview of the facility. 

N 

FIGURE 2-17. WICHITA STATE UNIVERSITY 7- x 10-ft WIND TUNNEL FACILITY 

2.2.2 Test Model and Installation. 

Details of the finite swept wing model used in the aerodynamic performance tests were shown in 

figures 2-2 and 2-3. An aileron control surface was added to the wing model for investigating 

the effect of ice accretions on the behavior of control surface effectiveness and hinge moments. 

The aileron had a 15-in. span starting at 42 inches from the wing root, and its leading edge sweep 

was 18.93o. The leading edge and hinge line were at 75% and 80% of the wing local chord 

respectively. Note that, due to the wing taper, the local wing chord varied linearly with spanwise 

distance. The aileron surface behind the hinge line had a geometric mean chord of 0.2118 ft and 

a planform area of 0.2647 ft2. 

2-11
The wing model was installed as a reflection plane model, as shown in figures 2-18 to 2-20. It 

was instrumented with 203 pressure ports distributed chordwise at five spanwise locations, 

corresponding to 15%, 30%, 50%, 68%, and 85% semispan, as shown in figure 2-20. The 

number of chordwise pressure ports for each spanwise station was 52 ports at 15% semispan 

(near the wing root), 49 ports at 30% semispan, 41 ports at 50% semispan (mid-semispan), 32 

ports at 68% semispan, and 29 ports at 85% semispan (near the wing tip). 

Streamlined 

Body 

MAC 

Tunnel Floor 

Tunnel Ceiling 

Aileron

Balance Virtual 

Center (Trunnion)

42.63 

All dimensions are in inches 

FIGURE 2-18. WING AND STREAMLINED BODY IN WSU 

TEST SECTION—SIDE VIEW 

Streamlined

 Body 

Gap Between

 Wing and 

Streamlined Body Aileron 

Streamlined Body

Pressure Taps at 

15%, 30%, 50%, 

68% and 85% 

Semispan

85% 

68%

50%

30%

15%

FIGURE 2-19. INSTALLATION OF 

CLEAN WING IN WSU WIND 

TUNNEL—FRONT VIEW 

FIGURE 2-20. INSTALLATION OF 

CLEAN WING IN WSU WIND 

TUNNEL—SIDE VIEW 

2-12
A direct current servomotor was used for deflecting the aileron and a torque cell for measuring 

aileron hinge moments. The deflection of the aileron was measured with a linear variable 

differential transducer (LVDT). The servomotor, torque cell, and LVDT were housed inside the 

wing mount, which connected the wing to the external balance, as shown in figures 2-21 and 

2-22. 

Pressure 

Tubings 

Wing Rotation 

Mechanism 

Mount that 

Connects Wing 

to External 

B l

Torque Cell 

Servomotor 

LVDT, Measures 

Deflection Angle

Zoom-in of

housing, 

shown in 

Fig. 2.22 

LVDT

Torque Cell

Servomotor

FIGURE 2-21. INSTALLATION OF 

CLEAN WING ON TUNNEL 

TURNTABLE—REAR VIEW 

FIGURE 2-22. HOUSING FOR 

SERVOMOTOR, TORQUE CELL, AND 

LVDT (Photographed under the wing, near the 

trailing edge.) 

Two large aluminum (Al) brackets were designed to connect the wing to the tunnel external 

balance and to raise the wing model 4.4 inches above the tunnel floor. Raising the model above 

the tunnel floor was necessary to keep the wing root section above the floor boundary layer, 

which was approximately 1.5 inches high. Since the Al brackets were exposed to the flow, the 

wing aerodynamic measurements would have been affected by the aerodynamic loads on the 

brackets. To eliminate this problem, a wooden fuselage-like body was used to shield the wing 

support brackets from the airflow. The streamlined body was 4.436 inches high and 55.6 inches 

long. Aerodynamic forces on the body did not affect the wing force and moment measurements 

since it was not connected to the tunnel balance. Extensive computational flow dynamics (CFD) 

studies were performed prior to the experimental investigation to ensure that the streamlined 

body did not modify the wing flow field and that the flow over the body remained attached even 

at high angles of attack (beyond wing stall). A small 0.25-in. gap was allowed between the wing 

root section and the streamlined body to prevent contact between the wing and the body. The 

0.25-in. gap was sealed with a latex strip called dental dam, as shown in figures 2-23 and 2-24, to 

prevent flow leakage through the gap that could have modified the flow near the wing root. 

2-13
Dental Dam

Streamlined Body

Dental Dam 

FIGURE 2-23. CLOSE-UP OF WING 

ROOT AND DENTAL 

DAM—PRESSURE SURFACE 

FIGURE 2-24. CLOSE-UP OF WING 

ROOT AND DENTAL 

DAM—SUCTION SURFACE 

2.2.3 Icing Research Tunnel Ice Castings and LEWICE Ice Shapes. 

The ice shapes selected for the aerodynamic experiments included six polyurethane castings of 

actual ice accretions and seven simulated ice shapes defined with the NASA Glenn LEWICE 2.0 

ice accretion code. Icing conditions and ice shape notations are provided in tables 2-1 to 2-3. To 

assess the effects of ice roughness on aerodynamic performance, the LEWICE ice shapes were 

tested with and without simulated roughness. A total of 20 ice shapes were investigated during 

the experimental study. 

2.2.3.1 IRT Ice Shape Castings. 

The ice shapes tested were polyurethane castings of actual ice accretions obtained at the NASA 

Glenn IRT facility with the GLC-305 wing model, as summarized in tables 2-1 and 2-2. For 

each icing condition, a set of three castings were made from the top, middle, and bottom RLEs, 

and these castings were glued together using epoxy to make a full span IRT ice shape casting. A 

total of six IRT ice castings were produced: IRT-CS10, IRT-IS10, IRT-SC5, IRT-CS2, IRTCS22, and IRT-IPSF22, corresponding to icing conditions 1 through 6, respectively. Figures 2-

25 to 2-32 show the installation of various IRT ice shape castings on the wing model. Also, 

figure 2-33 presents close-up views of all six IRT ice shape castings. 

2-14
TABLE 2-3. ICING CONDITIONS FOR LEWICE ICE SHAPES 

Icing 

Condition Description 

AOA** 

(deg) 

V 

(mph)

Ttotal

(°F) 

LWC 

(g/m3

) 

MVD 

(μm) 

τ

(min)

1 

(Glaze)* 

Complete Scallop Condition 

(ID: LS-CS10 and LR-CS10) 4 250.0 25.0 0.68 20.0 10.0 

2 

(Glaze)* 

Incomplete Scallop Condition 

(ID: LS-IS10 and LR-IS10) 4 150.0 25.0 0.65 20.0 10.0 

3 (Rime)* Scaled Condition from 2D tests 

(ID: LS-SC5 and LR-SC5) 6 201.3 11.7 0.51 14.5 5.0 

4 

(Glaze)* 

Complete Scallop Condition 

(ID: LS-CS2 and LR-CS2) 4 250.0 25.0 0.68 20.0 2.0 

5 

(Glaze)* 

Complete Scallop Condition 

(ID: LS-CS22N and LR-CS22N) 4 250.0 25.0 0.68 20.0 22.5 

5 

(Glaze)+

Complete Scallop Condition 

(ID: LS-CS22S and LR-CS22S) 4 250.0 25.0 0.68 20.0 22.5 

6 

(Glaze)* 

Failed Ice Protection 

(ID: LS-IPSF22 and LR-IPSF22) 4 150.0 27.0 0.46 20.0 22.5 

The methodology used to define the LEWICE ice shapes is discussed in section 2.2.3. 

Roughness was simulated with 36-size loose grit (average roughness height is approximately 0.0211 inch or 0.5356 

mm). 

*Computed using airfoil sections and the component of velocity normal to the wing leading edge (see section 2.2.3). 

+

Computed using streamwise velocity and airfoil sections (see section 2.2.3). 

**AOA = Angle of attack. 

Installation 

Bracket at 

Wing Tip

Aluminum Tape is used to 

Attach Ice Shape to 

Leading Edge 

Pressure Ports 

Installation 

Bracket at 

Wing Root

Gap between Ice 

Shape and Top of 

Streamlined Body Dental Dam

FIGURE 2-25. ICING RESEARCH 

TUNNEL CS10 ICE SHAPE CASTING 

INSTALLED ON WING—SIDE VIEW 

FIGURE 2-26. CLOSE-UP OF IRT-CS10 

ICE SHAPE CASTING AT WING ROOT 

2-15
Installation 

Bracket Near 

Mid-semispan

Epoxy Joining 

Castings Together 

Pressure 

Ports

FIGURE 2-27. CLOSE-UP OF 

INSTALLATION OF IRT-IS10 ICE 

SHAPE CASTING ON WING MODEL 

FIGURE 2-28. CLOSE-UP OF 

INSTALLATION OF IRT-CS22 ICE 

SHAPE CASTING ON WING MODEL 

FIGURE 2-29. INSTALLATION OF IRTSC5 ICE SHAPE CASTING ON 

WING MODEL—SIDE VIEW 

FIGURE 2-30. ICING RESEARCH 

TUNNEL CS2 ICE SHAPE CASTING ON 

TEST MODEL—FRONT VIEW 

2-16
FIGURE 2-31. GLC-305 SWEPT WING 

WITH IRT-CS22 ICE SHAPE CASTING—

FRONT VIEW 

FIGURE 2-32. FRONT VIEW OF IRTIPSF22 ICE SHAPE CASTING 

AT WING ROOT 

(a) View of IRT ice shapes near wing tip – Tip segments 

IRT-CS22 

IRT-CS10 

IRT-IS10 

IRT-SC5 

IRT-CS2 

IRT-IPSF22

IRT-CS22 IRT-IPSF22

IRTCS2

IRTSC5 IRT-IS10 

IRT-CS10

(b) View of IRT ice shapes near wing root – Root segments 

FIGURE 2-33. POLYURETHANE CASTINGS OF THE SIX ICE SHAPES OBTAINED 

FROM ICING TESTS AT IRT 

2-17
For ease of installation during testing, ice shape castings were made with cuffs that wrapped 

around the wing LE. The small ice shape castings were mounted to the wing using aluminum 

tape. The larger castings, IRT-CS10, IRT-CS22, and IRT-IPSF22, had aluminum brackets at 

selected spanwise locations that allowed these castings to be attached securely to the wing. This 

was necessary due to the considerable aerodynamic loads experienced by these large ice shapes. 

The leading-edge cuffs and the brackets on the IRT castings provided the additional benefit of 

consistent ice shape attachment to the wing during repeated installations. 

2.2.3.2 Smooth LEWICE Ice Shapes. 

The simulated shapes were 3D ice shapes that were defined from a series of 2D ice sections 

obtained with the LEWICE ice accretion code [5]. Table 2-3 summarizes all icing conditions for 

LEWICE analyses. The procedure used to define the 3D LEWICE ice shapes is detailed in 

reference 6 and is summarized below: 

1. Streamwise wing sections at 0% (wing root), 15%, 50%, 85%, and 100% (wing tip) 

semispan were selected for the development of the LEWICE ice shapes. 

2. Next, four additional wing sections were defined by taking the intersection of the wing 

with planes normal to the wing LE at 15%, 50%, 85%, and 100% semispan. The LE of 

the normal sections was at the same spanwise location as the streamwise sections defined 

in step 1. 

3. 3D Navier-Stokes computations were performed at WSU with the clean wing. In the 

computations, the IRT walls were included to simulate the tunnel wall effects on the wing 

flow field. The geometric angles of attack (α) used in the analysis were 4° and 6° to 

match the angles of attack in the icing tests. Analysis pressures for streamwise sections 

at 15%, 50%, and 85% semispan were compared with experimental pressure distributions 

obtained in the IRT facility. Good correlation between experiment and analysis was 

demonstrated. From the computed flow fields, pressure distributions were obtained for 

the four wing sections normal to the wing LE defined in step 2 and for the streamwise 

section at the wing root. 

4. 2D ice accretion analyses were conducted with the LEWICE 2.0 computer code. The 

computations were performed using five sections of the GLC-305 swept finite wing. 

These included the streamwise section at the wing root and the four sections normal to 

the wing LE at 15%, 50%, 85%, and 100% semispan (see step 2 above). The icing 

conditions for the LEWICE analysis were identical to those used in the IRT icing tests 

(see tables 2-1 and 2-3). However, the LEWICE angle of attack for each wing section 

was adjusted to match the pressure distributions from the Navier-Stokes analysis 

described in step 3. The velocity for the ice accretion computations conducted with the 

four sections defined normal to the wing leading edge was set to equal the component of 

the free-stream normal to the wing LE. 

5. For each icing condition in table 2-3, five LEWICE ice shape sections (one streamwise 

and four normal to the wing leading edge) were obtained. The five ice shape sections 

2-18
along with the wing geometry were imported into a computer-aided design software 

package and were used to define 3D LEWICE ice shapes. The 2D ice shape sections at 

the five spanwise locations were connected with spline surfaces and with plane surfaces. 

A comparison of the 3D ice shapes obtained using spline interpolation with 

corresponding ice shapes obtained with linear interpolation showed very small 

differences. Since the spline interpolation method provided a smooth transition in the 

spanwise direction between the five 2D sections, it was used to define the 3D LEWICE 

ice shapes, corresponding to icing conditions 1 to 4 (ice shapes LS-CS10, LS-IS10, LSSC5, and LS-CS2). Straight-line interpolation was used for the remaining two ice shapes 

(LS-CS22N and LS-IPSF22), which were the large 22.5-min glaze ice accretions. The 

reason for using straight-line interpolation for LS-CS22N and LS-IPSF22 was to match 

the method used by some aircraft manufacturers in defining 3D LEWICE ice shapes from 

2D sections. 

The use of airfoil sections normal to the wing LE along with the normal component of the 

velocity vector for the LEWICE analysis was based on input from NASA. However, some 

airframers prefer to use streamwise airfoil sections and the streamwise velocity in defining 3D 

ice shapes with the LEWICE computer code. Typically, streamwise ice accretion analyses 

produce ice shapes with larger horns (for glaze ice shapes) due to the higher water load resulting 

from the higher speed (streamwise velocity is greater than the velocity component normal to the 

LE). To compare the aerodynamic effects of LEWICE ice shapes based on normal and 

streamwise wing sections and flow velocities, one more (seventh) LEWICE ice shape was 

defined. This shape was obtained using the procedure discussed in steps 3 to 5 above. However, 

for this ice shape, streamwise wing sections at 0%, 15%, 50%, 85%, and 100% semispan were 

used along with the streamwise free-stream speed to define five 2D LEWICE ice shapes. The 

icing condition used for the seventh LEWICE ice shape (LS-CS22S) was icing condition 5 in 

table 2-1, which in the IRT icing tests produced a 22.5-min scalloped ice shape with large horns. 

Close-up views of all LEWICE ice shapes are shown in figure 2-34. The smooth LEWICE 

sections and corresponding IRT ice shape tracings are compared in figures 2-35 to 2-40. Section 

comparisons were made at spanwise locations labeled cut A, cut B, and cut C in figure 2-5. The 

sections at stations A and B were taken normal to the wing leading edge. The section at station 

C was in the streamwise direction. Section C was at the wing root, section B was 25 inches from 

the root (measured along the wing leading edge), and section A was at 50 inches from the root 

(measured along the wing leading edge). Note that the distance from the wing root to the wing 

tip measured along the wing leading edge was 67.95 inches. Figures 2-35 to 2-40 demonstrate 

that the ice accretions had a notable twist from the tip of the wing towards the root. This was 

due to the geometric twist of the wing. For all ice shapes obtained at α of 4°, the wing tip 

section was at αlocal of 0°, while the wing root was at +4°. Thus, the ice on the root accreted 

towards the lower surface of the wing; while at the wing tip, it was almost equally distributed on 

both surfaces. In comparing the LEWICE sections with the tracings of the IRT ice shapes, it is 

important to realize that the traces do not reflect the complex 3D features of the IRT ice shapes. 

A comparison between the 3D LEWICE and IRT ice casting shapes for icing condition 2 is 

provided in figure 2-41. Figures 2-42 and 2-43 compare the profiles of LEWICE ice shapes 

obtained for icing condition 5 using streamwise and normal sections to the wing LE as discussed 

above. 

2-19
The LS-CS10, LS-IS10, LS-SC5, LS-CS2, and LS-IPSF22 ice shapes were made of aluminum, 

whereas the LS-CS22N and LS-CS22S ice shapes were fabricated of wood. All LEWICE ice 

shapes were manufactured with LE cuffs to secure and ensure proper installation onto the wing. 

The large ice shapes LS-CS10, LS-CS22N, LS-CS22S, and LS-IPSF22 had aluminum brackets 

to provide additional support and consistency during repeated installations. 

a. Wing Tip 

 LS-IPSF22 LS-CS22S LR-CS22N LR-CS2 LR-SC5 LR-IS10 LR-CS10 

b. Wing Root 

 LR-CS10 LR-IS10 LR-SC5 LR-CS2 LR-CS22N LS-CS22S LS-IPSF22 

FIGURE 2-34. SMOOTH AND ROUGH LEWICE ICE SHAPES 

2-20
-4.0 -3.0 -2.0 -1.0 0.0 1.0 2.0

X (in)

-3.0

-2.0

-1.0

0.0

1.0

2.0

3.0

Z (in)

IRT casting A LEWICE

-4.0 -3.0 -2.0 -1.0 0.0 1.0 2.0

X (in)

-3.0

-2.0

-1.0

0.0

1.0

2.0

3.0

Z (in)

IRT casting A LEWICE

(a) Profiles at station A (50 inches from root) (a) Profiles at station A (50 inches from root) 

-4.0 -3.0 -2.0 -1.0 0.0 1.0 2.0

X (in)

-3.0

-2.0

-1.0

0.0

1.0

2.0

3.0

Z (in)

IRT casting

LEWICE

B

-4.0 -3.0 -2.0 -1.0 0.0 1.0 2.0

X (in)

-3.0

-2.0

-1.0

0.0

1.0

2.0

3.0

Z (in)

IRT casting

LEWICE

B 

(b) Profiles at station B (25 inches from root) (b) Profiles at station B (25 inches from root) 

-4.0 -3.0 -2.0 -1.0 0.0 1.0 2.0

X (in)

-3.0

-2.0

-1.0

0.0

1.0

2.0

3.0

Z (in)

IRT casting

LEWICE

C

-4.0 -3.0 -2.0 -1.0 0.0 1.0 2.0

X (in)

-3.0

-2.0

-1.0

0.0

1.0

2.0

3.0

Z (in)

IRT casting

LEWICE

C 

(c) Profiles at station C (wing root) (c) Profiles at station C (wing root) 

FIGURE 2-35. COMPARISON OF IRT 

AND LEWICE ICE SHAPES FOR ICING 

CONDITION 1 (CS10) 

FIGURE 2-36. COMPARISON OF IRT 

AND LEWICE ICE SHAPES FOR ICING 

CONDITION 2 (IS10) 

2-21
-4.0 -3.0 -2.0 -1.0 0.0 1.0 2.0

X (in)

-3.0

-2.0

-1.0

0.0

1.0

2.0

3.0

Z (in)

IRT casting A LEWICE

-4.0 -3.0 -2.0 -1.0 0.0 1.0 2.0

X (in)

-3.0

-2.0

-1.0

0.0

1.0

2.0

3.0

Z (in)

IRT casting A LEWICE

(a) Profiles at station A (50 inches from root) (a) Profiles at station A (50 inches from root) 

-4.0 -3.0 -2.0 -1.0 0.0 1.0 2.0

X (in)

-3.0

-2.0

-1.0

0.0

1.0

2.0

3.0

Z (in)

IRT casting

LEWICE

B

-4.0 -3.0 -2.0 -1.0 0.0 1.0 2.0

X (in)

-3.0

-2.0

-1.0

0.0

1.0

2.0

3.0

Z (in)

IRT casting

LEWICE

B 

(b) Profiles at station B (25 inches from root) (b) Profiles at station B (25 inches from root) 

-4.0 -3.0 -2.0 -1.0 0.0 1.0 2.0

X (in)

-3.0

-2.0

-1.0

0.0

1.0

2.0

3.0

Z (in)

IRT casting

LEWICE

C

-4.0 -3.0 -2.0 -1.0 0.0 1.0 2.0

X (in)

-3.0

-2.0

-1.0

0.0

1.0

2.0

3.0

Z (in)

IRT casting

LEWICE

C 

(c) Profiles at station C (wing root) (c) Profiles at station C (wing root) 

FIGURE 2-38. COMPARISON OF IRT 

AND LEWICE ICE SHAPES FOR ICING 

CONDITION 4 (CS2) 

FIGURE 2-37. COMPARISON OF IRT 

AND LEWICE ICE SHAPES FOR ICING 

CONDITION 3 (SC5) 

2-22
-4.0 -3.0 -2.0 -1.0 0.0 1.0 2.0

X (in)

-3.0

-2.0

-1.0

0.0

1.0

2.0

3.0

Z (in)

IRT casting A LEWICE

(a) Profiles at station A (50 inches from root) (a) Profiles at station A (50 inches from root) 

(b) Profiles at station B (25 inches from root) (b) Profiles at station B (25 inches from root) 

-4.0 -3.0 -2.0 -1.0 0.0 1.0 2.0

X (in)

-3.0

-2.0

-1.0

0.0

1.0

2.0

3.0

Z (in)

IRT casting

LEWICE (Normal)

B LEWICE (Streamwise)

A

-4.0 -3.0 -2.0 -1.0 0.0 1.0 2.0

X (in)

-3.0

-2.0

-1.0

0.0

1.0

2.0

3.0

Z (in)

IRT casting

LEWICE (Normal)

LEWICE (Streamwise)

-4.0 -3.0 -2.0 -1.0 0.0 1.0 2.0

X (in)

-3.0

-2.0

-1.0

0.0

1.0

2.0

3.0

Z (in)

IRT casting

LEWICE

C 

-4.0 -3.0 -2.0 -1.0 0.0 1.0 2.0

X (in)

-3.0

-2.0

-1.0

0.0

1.0

2.0

3.0

Z (in)

IRT casting

LEWICE

B

-4.0 -3.0 -2.0 -1.0 0.0 1.0 2.0

X (in)

-3.0

-2.0

-1.0

0.0

1.0

2.0

3.0

Z (in)

IRT casting

LEWICE

C

(c) Profiles at station C (wing root) (c) Profiles at station C (wing root) 

FIGURE 2-40. COMPARISON OF IRT 

AND LEWICE ICE SHAPES FOR ICING 

CONDITION 6 (IPSF22) 

FIGURE 2-39. COMPARISON OF IRT 

AND LEWICE ICE SHAPES FOR ICING 

CONDITION 5 (CS22) 

2-23
FIGURE 2-41. DETAILS OF ICE SHAPES OBTAINED FROM ICING CONDITION 2 

a. LS-IS10 Ice 

Shape—Root 

Section 

b. Comparison of LR-IS10 

and IRT-IS10 Ice 

Shapes—Root Section

c. Comparison of LR-IS10 

and IRT-IS10 Ice 

Shapes—Tip Section

a. LR-CS22N b. LR-CS22S

FIGURE 2-42. COMPARISON OF PROFILE OF LR-CS22N AND LR-CS22S 

ICE SHAPES—TIP SECTION 

a. LR-CS22N 

b. LR-CS22S 

FIGURE 2-43. COMPARISON OF PROFILE OF LR-CS22N AND LR-CS22S 

ICE SHAPES—ROOT SECTION 

2-24
2.2.3.3 Rough LEWICE Ice Shapes. 

The effect of ice shape roughness was simulated by adding 36-size (average roughness height is 

approximately 0.0211 inches or 0.5356 mm) loose grit to the smooth LEWICE ice shapes. The 

roughened LEWICE (LR) ice shapes were named LR-CS10, LR-IS10, LR-SC5, LR-CS2, LRCS22N, LR-CS22S, and LR-IPSF22. Figures 2-44 to 2-51 show the installation of LEWICE ice 

shapes on the wing model. 

Pressure Ports

FIGURE 2-45. CLOSE-UP OF LR-IS10 ICE 

SHAPE AT TIP OF WING MODEL 

FIGURE 2-44. INSTALLATION OF LRCS10 ROUGH LEWICE ICE SHAPE ON 

WING MODEL 

2-25
FIGURE 2-46. INSTALLATION OF LRSC5 ROUGH LEWICE ICE SHAPE ON 

WING MODEL 

FIGURE 2-47. INSTALLATION OF LRCS2 ROUGH LEWICE ICE SHAPE ON 

WING MODEL 

FIGURE 2-48. INSTALLATION OF 

LS-CS22N SMOOTH LEWICE ICE SHAPE 

ON WING MODEL 

FIGURE 2-49. INSTALLATION OF 

LR-CS22N ROUGH LEWICE ICE SHAPE 

ON WING MODEL 

2-26
FIGURE 2-51. INSTALLATION OF 

LS-IPSF22 ICE SHAPE ON WING 

MODEL—PRESSURE SIDE 

FIGURE 2-50. INSTALLATION OF 

LS-IPSF22 ICE SHAPE ON WING 

MODEL—SUCTION SIDE 

2.2.4 Test Procedure. 

The wind tunnel investigation was divided into four parts that included model installation and 

verification of all instrumentation, preliminary, production, and flow visualization tests. 

A significant effort was directed at the start of the experimental investigation to ensure that all 

necessary hardware and software for the aerodynamic experiments were in proper working order. 

Part of this effort included a complete external balance calibration, testing of the pressure 

instrumentation, and calibration of the aileron torque cell and LVDT. Next, a number of 

preliminary wind tunnel tests were conducted with the clean wing to evaluate test repeatability, 

assess tunnel flow angularity, and establish data-sampling rates and sampling period. Moreover, 

tests were performed to check force, moment, hinge moment, and pressure instrumentation, and 

to evaluate the method for mounting the ice shapes onto the wing LE. During the preliminary 

tests, the software for acquiring and reducing experimental data were tested for accuracy using 

known static loads that were applied to the wing and the aileron. 

The third segment of the investigation was the production runs. The first set of production runs 

were conducted with the clean wing using fixed and free transition. Next, tests were performed 

with all 20 ice shape configurations, which included six IRT ice shape castings, seven smooth 

LEWICE ice shapes, and seven rough LEWICE ice shapes. Conditions for the experimental 

investigation included AOA sweeps for ten aileron deflections, as shown in table 2-4. The 

Reynolds number (Re) for all the tests was fixed at 1.8 million, and it was computed using the 

wing MAC. For all runs, the complete set (6 components) of forces and moments for the wing 

were obtained. Aileron hinge moments and wing surface pressures were also recorded. A 

2-27
limited Reynolds number study was also conducted with the clean and selected iced wing 

configurations for ReMAC of 0.5, 1.0, 1.5, and 1.8 million. A complete run log for all the 

production runs is listed in appendix B. 

TABLE 2-4. TEST CONDITIONS FOR AERODYNAMIC PERFORMANCE DATA 

Reynolds number based on wing MAC (million) 1.8 

Dynamic pressure, Q (psf) 50 

Free-stream Mach number, M∞ 0.185 

Free-stream velocity (V∞) 225.3 ft/s, 68.7 m/s, 153.7 mph, 133.5 kts 

Mach number normal to wing LE 0.16 

Wing geometric α-range (deg.) -8 to +20 by 1°

Aileron deflection, δA (deg.) -15, -10, -5, -2.5, 0, 2.5, 5, 10, 15, 20 

Static tares were obtained for the clean wing and for all ice shapes investigated. There was no 

need to obtain dynamic tares for the wing model since the model mount was not exposed to the 

flow, and the streamline body fairing was not connected to the external balance. 

Flow angularity and wing downwash were measured for a portion of the runs. This was 

accomplished with two, seven-hole flow angularity probes, designed and calibrated by 

AEROPROBE. The seven-hole probes were installed on the clean wing at two stations 

corresponding to approximately 16% and 80% semispan. Seven-hole probes were selected 

because they can measure the three components of velocity, the total pressure, and the static 

pressure at a point in the flow. The seven-hole flow probes provided results with high accuracy 

for flow angles as high as 75° [7]. The data from the seven-hole probes were used to estimate 

local α near the inboard and outboard sections of the wing. The tests with the flow probes 

included α-sweeps for three aileron deflections (0°, -15°, and +20°). The Reynolds number in all 

cases was 1.8 million, based on wing MAC. Installation of the seven-hole flow probes is 

depicted in figures 2-52 to 2-55. 

At the end of production runs, flow visualization tests were performed for selected 

configurations, which included all IRT castings, two smooth LEWICE ice shapes (LS-CS22S 

and LS-IPSF22), and all rough LEWICE ice shapes. Visualization of the clean and iced wing 

flow fields was accomplished using white yarn tufts attached to both surfaces of the wing model 

and the wing fairing. Each flow visualization run was performed at a Reynolds number of 

1.5 million based on wing MAC and consisted of an α-sweep (-8° to +16°, increment of 1°) with 

the aileron in the neutral position. Figures 2-56 and 2-57 show example installations of yarn 

tufts on the clean and iced wing. Three video cameras were used to monitor the flow pattern on 

both surfaces of the model. One of the cameras, which was positioned outside the south wall of 

the tunnel, was used to capture the flow pattern on the suction (upper) surface of the wing. The 

other two cameras were located above the window at the tunnel ceiling and outside the tunnel’s 

north wall. These cameras recorded the flow pattern on the pressure side of the wing. Separated 

flow patterns near the wing tip of an iced wing configuration are illustrated in figures 2-58 and 

2-59. 

2-28
Flow Probe Near to 

Wing Tip 

Flow Probe Near 

to Wing Root 

Flow Probe Near to 

Wing Root 

Flow Probe Near to 

Wing Tip 

FIGURE 2-52. INSTALLATION OF 

SEVEN-HOLE FLOW PROBES ON 

WING MODEL—FRONT VIEW 

FIGURE 2-53. INSTALLATION OF 

SEVEN-HOLE FLOW PROBES ON 

WING MODEL—SIDE VIEW 

Flow Probe 

Installation 

Bracket 

Flow Probe 

Installation 

Bracket 

FIGURE 2-55. CLOSE-UP OF SEVENHOLE FLOW PROBE NEAR TO WING TIP 

(80% SEMISPAN) 

FIGURE 2-54. CLOSE-UP OF SEVENHOLE FLOW PROBE NEAR TO WING 

ROOT (16% SEMISPAN) 

2-29
FIGURE 2-57. INSTALLATION OF 

YARN TUFTS ON ICED WING 

AND STREAMLINED 

BODY—SUCTION SURFACE 

FIGURE 2-56. TUFTS INSTALLATION 

ON CLEAN WING AND STREAMLINED 

BODY—PRESSURE SURFACE 

FIGURE 2-58. TUFTS SHOW THAT 

FLOW DOWNSTREAM OF ICE SHAPE 

WAS SEPARATED NEAR WING TIP 

FIGURE 2-59. TUFTS SHOW COMPLETE 

FLOW SEPARATION AT WING TIP 

2-30
2.2.5 Test Measurements. 

Aerodynamic forces and moments were measured with a six-component truncated pyramid-type 

external balance. Aileron hinge moments were measured with a 10-in-lb temperature 

compensated torque sensor. The resolution of the hinge-moment torque cell was 0.005 in-lb. 

Force balance measurements were resolved as follows: ±0.75 lb for lift, ±0.3 lb for drag, and 

±0.2 ft-lb for pitching moment. The resolution of the aerodynamic coefficients obtained at a 

tunnel dynamic pressure of 50 psf was 0.002 for lift coefficient (CL), 0.0008 for drag coefficient 

(CD), 0.0003 for pitching-moment coefficient (CM), and 0.0001 for hinge-moment coefficient 

(CH). 

Pressure measurements were obtained with a Pressure Systems Inc. 8400 Industrial System 

processor. Pressures were resolved to ±0.0025 psi corresponding to a pressure coefficient (Cp) 

of 0.0072 for a test dynamic pressure of 50 psf. 

Aerodynamic coefficients for each α were computed by taking the average of two sets of 

measurements. Each set consisted of 1024 data points (per channel) obtained over a time 

interval of 4 seconds. Thus, the value of each of the force and moment coefficients was 

computed by taking the average of 2048 data points per α. 

2-31/2-32
3. RESULTS AND DISCUSSION. 

This section discusses the characteristics of the six ice accretions obtained from the icing tests at 

the NASA IRT facility and selected experimental data from aerodynamic experiments conducted 

at the WSU wind tunnel facility. Additional information regarding the icing tests is given in 

reference 8. 

3.1 Icing Conditions and Ice Accretions (NASA IRT). 

Conditions for the IRT icing tests are summarized in table 2-1 and include five glaze conditions 

with ice accretion times (τ) of 2, 10, and 22.5 minutes and a 5-min rime ice case. Ice accretion 

characteristics for each icing condition are discussed below. 

3.1.1 Icing Condition 1—Complete Scallop Condition (τ = 10 min). 

At this glaze icing condition, the ice accretion on the top RLE segment (figure 3-1) shows welldefined scallop tips beginning at the attachment lines. The space between the scallop tips shows 

some accumulation of ice. The scallop tips are covered with roughness elements. The side view 

of the ice accretion on the pressure side of the wing (lower surface), shown in figure 3-2, 

demonstrates the feather formation of the scallop tips. Figure 3-3 provides an overall view of the 

ice accretion obtained at the top RLE. The casting photograph in figure 3-4 shows how the ice 

accretion increases in size away from the tip of the wing. The ice accretions on the middle and 

bottom RLEs exhibited the same characteristics as the ones observed with the ice on the top 

RLE. 

FIGURE 3-1. ICING CONDITION 1, 

FRONT VIEW OF ICE ACCRETION ON 

THE TOP RLE 

FIGURE 3-2. ICING CONDITION 1, SIDE 

VIEW OF ICE ACCRETION ON 

PRESSURE SIDE OF THE TOP RLE 

3-1
FIGURE 3-3. ICING CONDITION 1, 

FRONT VIEW OF THE COMPLETE ICE 

ACCRETION ON THE TOP RLE 

FIGURE 3-4. ICING CONDITION 1, 

CASTING FROM THE ICE ACCRETION 

ON THE TOP RLE 

(a) Front View

(b) Side 

View 

3.1.2 Icing Condition 2—Incomplete Scallop Condition (τ = 10 min). 

For this glaze icing condition, which is typical of an incomplete scallop case, the ice shape on the 

bottom RLE, shown in figure 3-5, exhibits scallop tips beginning at some distance from the 

attachment line. The side view of the ice accretion on the suction side of the bottom RLE 

segment is presented in figure 3-6 and shows the scallop tips and their feather structure. The ice 

accretions on the top and middle leading-edge segments exhibited the same characteristics. 

Figure 3-7 presents an overall view of the ice accretion on the bottom RLE, while figure 3-8 

shows front and side views of the ice casting for this wing segment. The side view in figure 3-8 

demonstrates the ice shape characteristics on the pressure side of the bottom RLE segment of the 

wing. 

3.1.3 Icing Condition 3—Scaled Condition (τ = 5 min). 

This condition was obtained by scaling a reference icing condition [9] tested with a 2D 36-in. 

chord GLC-305 wing model. The icing condition for the 2D wing was α = 6°, V∞ = 201.3 mph, 

Ttotal = 11.7°F, LWC = 0.40 g/m3

, MVD = 20 μm, and τ = 16.7 minutes. Aerodynamic tests with 

the 2D airfoil and the ice shape obtained with the reference icing condition were conducted in the 

Low-Turbulence Pressure Tunnel (LTPT) at NASA Langley Research Center [10]. Comparison 

of the aerodynamic performance data from the LTPT tests with corresponding results from the 

WSU finite wing tests using the scaled ice shape will help evaluate differences between 2D and 

3D iced wing aerodynamics. 

3-2
FIGURE 3-5. ICING CONDITION 2, 

FRONT VIEW OF ICE ACCRETION ON 

THE BOTTOM RLE 

FIGURE 3-6. ICING CONDITION 2, SIDE 

VIEW OF ICE ACCRETION ON SUCTION 

SIDE OF THE BOTTOM RLE 

FIGURE 3-7. ICING CONDITION 2, 

FRONT VIEW OF THE COMPLETE ICE 

ACCRETION ON THE BOTTOM RLE 

FIGURE 3-8. ICING CONDITION 2, 

CASTING FROM THE ICE ACCRETION 

ON THE BOTTOM RLE 

(a) Front 

View 

(b) Side 

View 

The ice accretion presented in figure 3-9 is a no-scallop case. The side view of the ice accretion 

on the pressure side of the wing, depicted in figure 3-10, shows a feather structure. The ice 

around the attachment line is smooth with a pointed shape and a whitish color. This indicates 

that the ice accretion is of the rime type. Figure 3-11 shows an overall view of the ice accretion 

3-3
obtained on the bottom RLE. Figure 3-12 shows a front and side view of the casting from the ice 

accretion on the bottom RLE. The side view corresponds to the suction side of the wing. The 

ice on the top, middle, and bottom RLEs show the same characteristics. 

FIGURE 3-9. ICING CONDITION 3, 

FRONT VIEW OF ICE ACCRETION ON 

THE BOTTOM RLE 

FIGURE 3-10. ICING CONDITION 3, SIDE 

VIEW OF ICE ACCRETION ON SUCTION 

SIDE OF THE BOTTOM RLE 

(a) Front 

View 

(b) Side 

View

FIGURE 3-11. ICING CONDITION 3, 

FRONT VIEW OF THE COMPLETE ICE 

ACCRETION ON THE BOTTOM RLE 

FIGURE 3-12. ICING CONDITION 3, 

CASTING FROM THE ICE ACCRETION 

ON THE BOTTOM RLE 

3-4
3.1.4 Icing Condition 4—Complete Scallop Condition (τ = 2 min). 

This icing condition was chosen to obtain an ice shape for a complete scallop case at a short ice 

accretion time. The only difference between this condition and icing conditions 1 and 5 was the 

duration of the ice accretion time: 2 minutes versus 10 and 22.5 minutes, respectively. The ice 

accretion in figure 3-13 shows the presence of scallop tip. The scallop tips on the suction side of 

the wing were more developed than on the pressure side. Figure 3-14 shows the ice accretion 

scallop tips and their feather formation on the suction side of the wing. Figure 3-15 provides an 

overall view of the ice accretion obtained over the middle RLE. Figure 3-16 shows the front and 

side views of the ice casting from the ice accretion on the middle RLE. The side view of the ice 

casting corresponds to the suction side of the wing. 

FIGURE 3-13. ICING CONDITION 4, 

FRONT VIEW OF ICE ACCRETION ON 

THE MIDDLE RLE 

FIGURE 3-14. ICING CONDITION 4, SIDE 

VIEW OF ICE ACCRETION ON SUCTION 

SIDE OF THE MIDDLE RLE 

(a) Front 

View (b) Side 

View 

FIGURE 3-15. ICING CONDITION 4, 

FRONT VIEW OF THE COMPLETE ICE 

ACCRETION ON THE MIDDLE RLE 

FIGURE 3-16. ICING CONDITION 4, 

CASTING FROM THE ICE ACCRETION 

ON THE MIDDLE RLE 

3-5
3.1.5 Icing Condition 5—Complete Scallop Condition (τ = 22.5 min). 

This icing case was selected to obtain a complete scallop ice shape for a long ice accretion time. 

In general, the ice shape in figure 3-17 had similar characteristics as icing condition 1 but its size 

was considerably larger. Figure 3-18 shows a side view of the ice accretion on the suction side 

of the wing that demonstrates the ice shape scallop tips and their feather structure. Figure 3-19 

shows an overall view of the ice accretion obtained over the middle RLE. Figure 3-20 provides 

front and side views of the casting from the ice accretion on the middle RLE. The side view is of 

the wing pressure side. 

FIGURE 3-17. ICING CONDITION 5, 

FRONT VIEW OF ICE ACCRETION ON 

THE MIDDLE RLE 

FIGURE 3-18. ICING CONDITION 5, SIDE 

VIEW OF ICE ACCRETION ON SUCTION 

SIDE OF THE MIDDLE RLE 

(a) Front View 

(b) Side 

View

FIGURE 3-19. ICING CONDITION 5, 

FRONT VIEW OF THE COMPLETE ICE 

ACCRETION ON THE MIDDLE RLE 

FIGURE 3-20. ICING CONDITION 5, 

CASTING FROM THE ICE ACCRETION 

ON THE MIDDLE RLE 

3-6
When this icing condition was run with the wing model in the high position, a piece of the ice 

accretion was consistently shed off the wing. For this reason, the velocity for icing condition 5 

was reduced to 225 mph with the wing in the high position. To maintain the same water mass 

flux as in the case at 250 mph, the LWC was increased to 0.89 g/m3

. Photographic data from the 

ice accretion at 250 mph and the one obtained at 225 mph indicated that the two ice accretions 

were very close in overall characteristics. The ice castings made from the ice accretion obtained 

at 225 mph and 0.89 g/m3

 were the middle and bottom RLEs. The ice casting for the top RLE 

segment was obtained from the ice accretion tests conducted at a speed of 250 mph and LWC of 

0.68 g/m3

. 

Icing conditions 1, 4, and 5 were selected to investigate the stages in the evolution of complete 

scallop ice shapes obtained with progressively longer time and the aerodynamic performance 

effects of these ice shapes. 

3.1.6 Icing Condition 6—IPS Failure Condition (τ = 22.5 min). 

This condition was chosen to be within the FAA Title 14 Code of Federal Regulation Part 25 

Appendix C envelopes [11] to provide an ice shape that was representative of an ice protection 

system failure case. The ice accretion depicted in figure 3-21 exhibited scallop tips, which were 

not well defined. The scallop formation appears to have started away from the attachment line, 

although this is not certain because glaze ice covered the area around the attachment line. In 

addition, the ice was not transparent enough to allow determination of the presence of feathers 

around the attachment line area. A side view of the ice accretion (figure 3-22) shows the scallop 

tips formed by the glaze ice feathers. Figure 3-23 provides an overall view of the ice accretion 

obtained on the top RLE. A significant accumulation of ice on the wing tip was observed, and in 

general, the ice accretion was not uniform, as shown in figure 2-33. Over the leading edges of 

the bottom and middle removable segments, the glaze ice accretion exhibited indistinct scallop 

tips, and glaze ice covered the attachment line area. Figure 3-24 shows front and side views of 

the casting obtained from the ice accretion on the top RLE. The side view corresponds to the 

wing suction side (upper surface). 

FIGURE 3-21. ICING CONDITION 6, 

FRONT VIEW OF ICE ACCRETION ON 

THE TOP RLE 

FIGURE 3-22. ICING CONDITION 6, 

SIDE VIEW OF ICE ACCRETION ON 

PRESSURE SIDE OF THE TOP RLE 

3-7
FIGURE 3-23. ICING CONDITION 6, 

FRONT VIEW OF THE COMPLETE ICE 

ACCRETION ON THE TOP RLE 

FIGURE 3-24. ICING CONDITION 6, 

CASTING FROM THE ICE ACCRETION 

ON THE TOP RLE 

(a) Front 

View 

(b) Side 

View 

3.2 AERODYNAMIC PERFORMANCE STUDIES. 

Experimental lift, drag, pitching-moment, hinge-moment, and surface pressure coefficients for 

the clean and iced wing of selected test cases are presented in this section. All aerodynamic 

coefficients have been corrected for tunnel wall effects using the procedure discussed in 

reference 12. Pressure coefficients were computed using the corrected tunnel dynamic pressure. 

No other corrections were applied to the pressure data. Note that in all cases, the angles of attack 

in the figures are the geometric angles of attack and have not been corrected for upwash or 

downwash effects. 

Pitching-moment measurements were resolved to the quarter-chord location of the mean 

aerodynamic chord. By convention, nose-up pitching moment is considered positive, as 

illustrated in figure 3-25. 

Positive CM,MAC/4 Positive CH

Positive δA

Pressure 

Surface 

Suction 

Surface

FIGURE 3-25. NOTATION FOR PITCHING MOMENT, HINGE MOMENT, AND 

AILERON DEFLECTION ANGLE 

Aileron hinge-moment coefficients were obtained from the following formula: 

aa

H cSV

Hm C ⋅⋅⋅⋅ =

∞

2 5.0 ρ (3-1)

3-8
where Hm is the hinge-moment and Sa and ca are the aileron planform area and the aileron mean 

chord behind the hinge line respectively. (The values for Sa and ca are provided in figure 2-3.) 

The sign convention for CH is positive for hinge moments causing the aileron trailing edge to 

move down. Aileron deflections corresponding to trailing edge down are also considered 

positive. Figure 3-25 depicts the sign convention for CH and δA. 

In the following discussion, the hinge-moment curve is divided into three regions (A, B, and C), 

corresponding to the linear, near-stall, and poststall ranges of the lift curve, as shown in 

figure 3-26. Typically, in region A, the decrease of CH was linear, small and gradual, while in 

region B, the growth was much greater due to increasing flow separation over the control 

surface. Note the anomaly or break from region A to region B with increase CL and the 

differences between the clean and iced configurations. In region C, the hinge moment remained 

nearly constant, since beyond stall the pressure distribution over the aileron upper surface did not 

vary significantly for positive α. 

(a) Clean configuration, CL > 0 (b) Clean configuration, CL < 0 

(c) Iced configuration, CL > 0 (d) Iced configuration, CL < 0 

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Lift coefficient

-0.12

-0.10

-0.08

-0.06

-0.04

-0.02

0.00

0.02

0.04

0.06

0.08

Hinge moment coefficient

CH 

CL 

B C 

A

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Lift coefficient

-0.12

-0.10

-0.08

-0.06

-0.04

-0.02

0.00

0.02

0.04

0.06

0.08

Hinge moment coefficient

C B A

CL 

CH 

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

Lift coefficient

-0.10

-0.08

-0.06

-0.04

-0.02

0.00

0.02

0.04

0.06

Hinge moment coefficient

CL 

CH 

B A

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

Lift coefficient

-0.10

-0.08

-0.06

-0.04

-0.02

0.00

0.02

0.04

0.06

Hinge moment coefficient

CH

CL

A B C 

FIGURE 3-26. REGIONS A, B, AND C FOR HINGE-MOMENT CURVE OF CLEAN AND 

ICED CONFIGURATIONS 

3-9
All pressure data are presented in the form of Cp versus chordwise distance (x/c). In all cases, 

negative Cp coefficients corresponding to the suction side (upper surface) of the wing are plotted 

upward. 

3.2.1 Data Quality and Repeatability. 

The WSU wind tunnel facility has been used for commercial and research testing over the years. 

Facility personnel have considerable expertise in wind tunnel testing, instrumentation, data 

acquisition, and data processing. To obtain quality data, careful planning and well thought out 

test were implemented during the aerodynamic investigation as discussed below. 

• An analysis was performed to determine instrumentation sensitivity needed to provide the 

required resolution of force, moment, and pressure coefficients for the primary test 

condition, which was α-sweeps at Reynolds number based on MAC of 1.8 million. 

• Prior to WSU wind tunnel entry, a complete external balance calibration was performed. 

Instrumentation such as pressure transducers, torque tubes, load cells, flow probes, and 

model hardware were checked to ensure proper operation and accuracy. 

• The repeatability of the experimental data is a function of instrumentation, data-sampling 

rates, flow quality, flow unsteadiness, and model setup procedures. To verify the 

repeatability of the experimental setup, tests with the baseline configuration were 

repeated, and the results were compared. Repeatability of lift, drag, pitching-moment, 

and hinge-moment coefficients are provided in figure 3-27. Also, figures 3-28 to 3-30 

show pressure distributions of 15%, 50%, and 85% semispan locations at α of -4°, 0°, 4°, 

8°, 12°, and 16°. The maximum average percentage difference of aerodynamic 

coefficients from individual test runs from the average of all test repeats was 1% for CL, 

2% for CD, 1% for CH, and 3% for Cp. Most of the variation occurring near wing stall 

was due to flow field unsteadiness and small amplitude model vibration. 

• At regular intervals during the wind tunnel tests, the balance, the hinge-moment system, 

and the pressure transducers were tested with known inputs to verify that they were 

working properly. For the balance tests, known forces were applied to the model, and 

force and moment data were obtained for a complete α-sweep at zero airspeed. 

Figures 3-31 and 3-32 show the setup for balance and hinge-moment system checking. 

• To ensure repeatability in the installation of the ice shape tested, all ice shapes were 

attached to cuffs designed to fit the wing LE. For the large ice shapes, brackets were also 

installed at selected spanwise locations to provide additional support and minimize 

deflections due to the aerodynamic loads. 

• Preliminary tests were conducted with all models prior to the start of the production runs 

to verify tunnel and model instrumentation and data acquisition hardware and software. 

Static tares were obtained for all ice shapes to account for the weight effects of each ice 

shape on the aerodynamic measurements. 

3-10
• A limited number of tests were conducted prior to the production runs to investigate datasampling periods, as shown in figure 3-33. All the force and moment coefficients 

obtained from wind tunnel tests are average values based on multiple measurements over 

a period of time. Sampling rate and sampling time depend on tunnel facility, test model, 

and airspeed, and are usually established experimentally for each tunnel facility. For the 

WSU tests, 1024 readings were taken per α for each force and moment coefficient. The 

sampling time was 4 seconds. For a dynamic pressure of 50 psf, the airspeed in the WSU 

wind tunnel was approximately 222 ft/s. Given that the MAC of the wing is 1.56 ft, 1 

second of data acquisition was equivalent to averaging the flow over the model 142 

times. In 1 second, the air had traversed a distance equal to 142 chord lengths. 

• All tests were conducted at constant Reynolds number, i.e., the speed of the tunnel was 

adjusted to maintain constant ReMAC. 

(a) Lift coefficient (b) Drag coefficient 

(c) Pitching-moment coefficient (d) Hinge-moment coefficient 

FIGURE 3-27. REPEATABILITY OF CL, CD, CM, AND CH; CLEAN CONFIGURATION; 

Re = 1.8x106; δA = 0°

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Lift coefficient

Clean (R006)

Clean (R195)

Clean (R376)

Clean (R573)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

0.00

0.05

0.10

0.15

0.20

0.25

0.30

Drag coefficient

Clean (R006)

Clean (R195)

Clean (R376)

Clean (R573)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.10

-0.08

-0.06

-0.04

-0.02

0.00

0.02

0.04

0.06

0.08

0.10

Pitching moment coefficient

Clean (R006)

Clean (R195)

Clean (R376)

Clean (R573)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.12

-0.10

-0.08

-0.06

-0.04

-0.02

0.00

0.02

0.04

0.06

0.08

Hinge moment coefficient

Clean (R006)

Clean (R195)

Clean (R376)

Clean (R573)

3-11
(a) Cp vs x/c (α = -4°) (b) Cp vs x/c (α = 0°) 

(c) Cp vs x/c (α = 4°) (d) Cp vs x/c (α = 8°) 

(e) Cp vs x/c (α = 12°) (f) Cp vs x/c (α = 16°) 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

Clean (R195)

Clean (R376)

Clean (R573)

Clean: min Cp = -2.3254 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

Clean (R195)

Clean (R376)

Clean (R573)

Clean: max Cp = 0.7279 at x/c = 0

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

Clean (R195)

Clean (R376)

Clean (R573)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

Clean (R195)

Clean (R376)

Clean (R573)

Clean: min Cp = -3.4081 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

Clean (R195)

Clean (R376)

Clean (R573)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

-2.5

-3.0

Pressure coefficient

Clean (R006)

Clean (R195)

Clean (R376)

Clean (R573)

FIGURE 3-28. REPEATABILITY OF PRESSURE DISTRIBUTIONS AT 15% SEMISPAN; 

CLEAN CONFIGURATION; Re = 1.8x106; δA = 0°

3-12
(a) Cp vs x/c (α = -4°) (b) Cp vs x/c (α = 0°) 

(c) Cp vs x/c (α = 4°) (d) Cp vs x/c (α = 8°) 

(e) Cp vs x/c (α = 12°) (f) Cp vs x/c (α = 16°) 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

Clean (R195)

Clean (R376)

Clean (R573)

Clean: min Cp = -3.1817 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

Clean (R195)

Clean (R376)

Clean (R573)

Clean: max Cp = 0.6843 at x/c = 0.002

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

Clean (R195)

Clean (R376)

Clean (R573)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

Clean (R195)

Clean (R376)

Clean (R573)

Clean: min Cp = -3.5221 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

Clean (R195)

Clean (R376)

Clean (R573)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

Clean (R195)

Clean (R376)

Clean (R573)

FIGURE 3-29. REPEATABILITY OF PRESSURE DISTRIBUTIONS AT 50% SEMISPAN; 

CLEAN CONFIGURATION; Re = 1.8x106; δA = 0°

3-13
(a) Cp vs x/c (α = -4°) (b) Cp vs x/c (α = 0°) 

(c) Cp vs x/c (α = 4°) (d) Cp vs x/c (α = 8°) 

(e) Cp vs x/c (α = 12°) (f) Cp vs x/c (α = 16°) 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

Clean (R195)

Clean (R376)

Clean (R573)

Clean: min Cp = -3.5329 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

Clean (R195)

Clean (R376)

Clean (R573)

Clean: max Cp = 0.6426 at x/c = 0.004

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

Clean (R195)

Clean (R376)

Clean (R573)

Clean: min Cp = -2.7034 at x/c = 0.004

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

Clean (R195)

Clean (R376)

Clean (R573)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

Clean (R195)

Clean (R376)

Clean (R573)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

Clean (R195)

Clean (R376)

Clean (R573)

FIGURE 3-30. REPEATABILITY OF PRESSURE DISTRIBUTIONS AT 85% SEMISPAN; 

CLEAN CONFIGURATION; Re = 1.8x106

; δA = 0°

3-14
(a) Front view of balance check setup (a) Rear view of hinge-moment load cell check setup 

(b) Rear view of balance check setup (b) Side view of hinge-moment load cell check setup 

(c) Close-up of weight (load) applied in normal to 

chord direction 

(c) Close-up of weight (load) applied in negative 

Hinge-moment direction (looking upstream) 

FIGURE 3-31. SETUP TO CHECK 

EXTERNAL BALANCE 

FIGURE 3-32. SETUP TO CHECK HINGEMOMENT LOAD CELL 

Load Applied 

Normal to Chord 

Direction 

Load Applied 

Parallel to Chord 

Direction 

3-15
(a) Lift coefficient (b) Drag coefficient 

(c) Pitching moment coefficient (d) Hinge-moment coefficient 

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Lift coefficient

FIGURE 3-33. DATA SAMPLING STUDIES OF CL, CD, CM, AND CH; CLEAN 

CONFIGURATION; Re = 1.8x106

; δA = 0°

In most wind tunnel studies, aerodynamic performance measurements rely on standard wind 

tunnel instrumentation, which typically include external or internal balances and, in some cases, 

pressure instrumentation. Near stall, extensive flow separation and vortex shedding increase 

model vibration and cause large wake unsteadiness, which can affect the accuracy of the 

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

1024 samples, 10 seconds

1024 samples, 8 seconds

1024 samples, 6 seconds

1024 samples, 4 seconds

1024 samples, 2 seconds

1024 samples, 1 second

2048 samples, 10 seconds

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

0.00

0.05

0.10

0.15

0.20

0.25

0.30

Drag coefficient

1024 samples, 10 seconds

1024 samples, 8 seconds

1024 samples, 6 seconds

1024 samples, 4 seconds

1024 samples, 2 seconds

1024 samples, 1 second

2048 samples, 10 seconds

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.10

-0.08

-0.06

-0.04

-0.02

0.00

0.02

0.04

0.06

0.08

0.10

Pitching moment coefficient

1024 samples, 10 seconds

1024 samples, 8 seconds

1024 samples, 6 seconds

1024 samples, 4 seconds

1024 samples, 2 seconds

1024 samples, 1 second

2048 samples, 10 seconds

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.10

-0.08

-0.06

-0.04

-0.02

0.00

0.02

0.04

0.06

0.08

Hinge moment coefficient

1024 samples, 10 seconds

1024 samples, 8 seconds

1024 samples, 6 seconds

1024 samples, 4 seconds

1024 samples, 2 seconds

1024 samples, 1 second

2048 samples, 10 seconds

3-16
measured aerodynamic coefficients. Considering the lift coefficient, the objective in most wind 

tunnel tests is to determine lift curve slope, near-stall and poststall behavior, and CL,max for 

engineering purposes. In determining the degree of measurement accuracy needed, an analysis 

should be performed based on the acceptable level of accuracy with respect to the aircraft 

performance. For example, assuming an error in CL of 0.01 for the wing used in this study, the 

corresponding error in lift at a speed of 131 kts (151 mph, Q = 50 psf) was approximately 3.7 lb. 

Given a CL,max of 0.87, the total lift generated by the wing at the same flow condition was 

319.7 lb. For engineering purposes, an error of 3.7 lb in lift out of a total lift of 319.7 lb should 

be considered to determine if better accuracy is needed in measuring CL,max. 

Precise measurement of aerodynamic properties near stall is a very difficult task, which is 

beyond the scope of the study described in this report. The experimental difficulties stem from a 

number of factors such as the ones listed below: 

• The exact dynamic responses of the balance with the model installed (e.g., response 

versus frequency) should be known and appropriate data-sampling rates and sampling 

periods should be established from experiments. The sampling periods should be long 

enough to account for both model and tunnel flow unsteadiness. 

• Model deflections and variations in α due to model vibration and flow pulsing are 

difficult to accurately measure or monitor in real time. 

• Tunnel blockage effects under stall conditions can be very significant, depending on 

model and tunnel size, and their impact on tunnel dynamic and static pressures are 

difficult to determine and correct. 

Generally, experimental data for unsteady conditions require special tunnel facilities and are 

usually obtained with simple geometries such as airfoil sections. In most cases, the best way to 

obtain such data is through the use of extensive surface pressure time histories, which can then 

be integrated to provide the required coefficients. 

In summary, the experimental data provided in this report have been obtained under carefully 

controlled conditions. Data near stall should be used with the understanding that it may be 

subject to some uncertainty. 

3.2.2 Compressibility Effects. 

For the clean GLC-305 airfoil section, the critical free-stream Mach number for 2D flow was 

determined to be 0.22. This value was obtained from the intersection of the two curves defined 

by equations 3-2 and 3-3 [13 and 14]. 

( )

⎥

⎥

⎥

⎦

⎤

⎢

⎢

⎢

⎣

⎡

− ⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

+

⋅−+

⋅ = − ∞

∞

1

1

122 1 2

2

γ

γ

γ

γ

γ

M

M Cpcr (3-2) 

3-17
0 2

2 2

2

0

12

2

1 1

1 Cp

M

M M

M

Cp Cp

∞

∞ ∞

∞

−

⎟

⎠

⎞ ⎜

⎝

⎛ ⋅ − +

+−

= γ

 (3-3) 

In the equations above, Cpcr is the critical pressure coefficient for which the local Mach number 

is 1, Cp0 is the maximum suction pressure coefficient for incompressible flow, and Cp is the 

compressible pressure coefficient obtained using Cp0 and Laitone’s compressibility correction 

given by equation 3-3. The value of Cp0 for the GLC-305 section was -10.1 and corresponded to 

α of 13.5°. This value was obtained from a 2D incompressible viscous flow analysis using the 

XFOIL [15] computer code. 

Note that the critical Mach number of 0.22 obtained from the 2D analysis and equations 3-2 and 

3-3 is a conservative estimate. Typically, in 3D flow, the value of the free-stream Mach number 

(M∞) required to achieve sonic flow over the wing for fixed α will be higher due to wing sweep 

and 3D flow relief effects. For the swept wing tested, the critical Mach number for 3D flow was 

0.249 ⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

28cos °

.220 . Since all tests at the WSU wind tunnel facility were conducted at M∞ of 0.185, 

the effect of compressibility on aerodynamic coefficients was not significant. 

3.2.3 Clean and Iced Wing Performance. 

Aerodynamic performance for the clean and iced wing is presented in figures 3-34 to 3-45 and in 

tables 3-1 to 3-9 for δA of 0°. The data presented demonstrates the effects of all the ice shapes 

(six IRT ice shapes, seven smooth LEWICE ice shapes, and seven rough LEWICE ice shapes) 

investigated on lift, drag, pitching moments, hinge moments, and on the surface pressure 

distributions. For the purpose of discussion, the percentages in tables 3-1, 3-2, 3-4, 3-5, 3-7, and 

3-8 have been calculated using the following formula: 

(%) property in Change ×100 − =Δ=

clean

iced clean

X

XX XX (3-4) 

and, increase in X means that a negative X becomes more negative and a positive X becomes 

more positive. 

3-18
(a) Lift coefficient (b) ΔCL vs α

(c) Drag coefficient (d) ΔCD vs α

(e) Pitching-moment coefficient (f) Hinge-moment coefficient 

FIGURE 3-34. EFFECT OF IRT ICE SHAPES ON CL, CD, CM, AND CH; IRT-CS10, 

IRT-IS10, IRT-SC5, IRT-CS2, IRT-CS22, AND IRT-IPSF22 CONFIGURATIONS; 

Re = 1.8x106

; δA = 0°

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Lift coefficient

Clean (R006)

IRT-CS10 (R018)

IRT-IS10 (R110)

IRT-SC5 (R055)

IRT-CS2 (R070)

IRT-CS22 (R047)

IRT-IPSF22 (R091)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-120

-100

-80

-60

-40

-20

0

20

40

ΔCL (%)

IRT-CS10 (R018)

IRT-IS10 (R110)

IRT-SC5 (R055)

IRT-CS2 (R070)

IRT-CS22 (R047)

IRT-IPSF22 (R091)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

0.00

0.05

0.10

0.15

0.20

0.25

0.30

0.35

0.40

0.45

Drag coefficient

Clean (R006)

IRT-CS10 (R018)

IRT-IS10 (R110)

IRT-SC5 (R055)

IRT-CS2 (R070)

IRT-CS22 (R047)

IRT-IPSF22 (R091)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-500

0

500

1000

1500

2000

2500

3000

3500

4000

ΔCD (%)

IRT-CS10 (R018)

IRT-IS10 (R110)

IRT-SC5 (R055)

IRT-CS2 (R070)

IRT-CS22 (R047)

IRT-IPSF22 (R091)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.10

-0.08

-0.06

-0.04

-0.02

0.00

0.02

0.04

0.06

0.08

0.10

Pitching moment coefficient

Clean (R006)

IRT-CS10 (R018)

IRT-IS10 (R110)

IRT-SC5 (R055)

IRT-CS2 (R070)

IRT-CS22 (R047)

IRT-IPSF22 (R091)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.12

-0.10

-0.08

-0.06

-0.04

-0.02

0.00

0.02

0.04

0.06

0.08

Hinge moment coefficient

Clean (R006)

IRT-CS10 (R018)

IRT-IS10 (R110)

IRT-SC5 (R055)

IRT-CS2 (R070)

IRT-CS22 (R047)

IRT-IPSF22 (R091)

3-19
(a) Cp vs x/c (α = -4°) (b) Cp vs x/c (α = 0°) 

(c) Cp vs x/c (α = 4°) (d) Cp vs x/c (α = 8°) 

(e) Cp vs x/c (α = 12°) (f) Cp vs x/c (α = 16°) 

FIGURE 3-35. EFFECT OF IRT ICE SHAPES ON PRESSURE DISTRIBUTIONS AT 15% 

SEMISPAN; IRT-CS10, IRT-IS10, IRT-SC5, IRT-CS2, IRT-CS22, AND IRT-IPSF22 

CONFIGURATIONS; Re = 1.8x106

; δA=0°

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

IRT-CS10 (R018)

IRT-IS10 (R110)

IRT-SC5 (R055)

IRT-CS2 (R070)

IRT-CS22 (R047)

IRT-IPSF22 (R091)

Clean: max Cp = 0.7099 at x/c = 0

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-CS10 (R018)

IRT-IS10 (R110)

IRT-SC5 (R055)

IRT-CS2 (R070)

IRT-CS22 (R047)

IRT-IPSF22 (R091)

Clean: min Cp = -2.3254 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

IRT-CS10 (R018)

IRT-IS10 (R110)

IRT-SC5 (R055)

IRT-CS2 (R070)

IRT-CS22 (R047)

IRT-IPSF22 (R091)

Clean: min Cp = -3.2976 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-CS10 (R018)

IRT-IS10 (R110)

IRT-SC5 (R055)

IRT-CS2 (R070)

IRT-CS22 (R047)

IRT-IPSF22 (R091)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

IRT-CS10 (R018)

IRT-IS10 (R110)

IRT-SC5 (R055)

IRT-CS2 (R070)

IRT-CS22 (R047)

IRT-IPSF22 (R091)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

-2.5

-3.0

Pressure coefficient

Clean (R006)

IRT-CS10 (R018)

IRT-IS10 (R110)

IRT-SC5 (R055)

IRT-CS2 (R070)

IRT-CS22 (R047)

IRT-IPSF22 (R091)

3-20
(a) Cp vs x/c (α = -4°) (b) Cp vs x/c (α = 0°) 

(c) Cp vs x/c (α = 4°)

(c) Cp vs x/c (α = 4°) (d) Cp vs x/c (α = 8°) 

(e) Cp vs x/c (α = 12°) (f) Cp vs x/c (α = 16°) 

FIGURE 3-36. EFFECT OF IRT ICE SHAPES ON PRESSURE DISTRIBUTIONS AT 50% 

SEMISPAN; IRT-CS10, IRT-IS10, IRT-SC5, IRT-CS2, IRT-CS22, AND IRT-IPSF22 

CONFIGURATIONS; Re = 1.8x106

; δA = 0°

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

IRT-CS10 (R018)

IRT-IS10 (R110)

IRT-SC5 (R055)

IRT-CS2 (R070)

IRT-CS22 (R047)

IRT-IPSF22 (R091)

Clean: max Cp = 0.677 at x/c = 0.002

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-CS10 (R018)

IRT-IS10 (R110)

IRT-SC5 (R055)

IRT-CS2 (R070)

IRT-CS22 (R047)

IRT-IPSF22 (R091)

Clean: min Cp = -3.1817 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

IRT-CS10 (R018)

IRT-IS10 (R110)

IRT-SC5 (R055)

IRT-CS2 (R070)

IRT-CS22 (R047)

IRT-IPSF22 (R091)

Clean: min Cp = -3.4743 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-CS10 (R018)

IRT-IS10 (R110)

IRT-SC5 (R055)

IRT-CS2 (R070)

IRT-CS22 (R047)

IRT-IPSF22 (R091)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

IRT-CS10 (R018)

IRT-IS10 (R110)

IRT-SC5 (R055)

IRT-CS2 (R070)

IRT-CS22 (R047)

IRT-IPSF22 (R091)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

IRT-CS10 (R018)

IRT-IS10 (R110)

IRT-SC5 (R055)

IRT-CS2 (R070)

IRT-CS22 (R047)

IRT-IPSF22 (R091)

3-21
(a) Cp vs x/c (α = -4°) (b) Cp vs x/c (α = 0°) 

(c) Cp vs x/c (α = 4°) (d) Cp vs x/c (α = 8°) 

(e) Cp vs x/c (α = 12°) (f) Cp vs x/c (α = 16°) 

FIGURE 3-37. EFFECT OF IRT ICE SHAPES ON PRESSURE DISTRIBUTIONS AT 85% 

SEMISPAN; IRT-CS10, IRT-IS10, IRT-SC5, IRT-CS2, IRT-CS22, AND IRT-IPSF22 

CONFIGURATIONS; Re = 1.8x106

; δA = 0°

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-CS10 (R018)

IRT-IS10 (R110)

IRT-SC5 (R055)

IRT-CS2 (R070)

IRT-CS22 (R047)

IRT-IPSF22 (R091)

Clean: min Cp = -3.5329 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-CS10 (R018)

IRT-IS10 (R110)

IRT-SC5 (R055)

IRT-CS2 (R070)

IRT-CS22 (R047)

IRT-IPSF22 (R091)

Clean: max Cp = 0.6244 at x/c = 0.004

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

IRT-CS10 (R018)

IRT-IS10 (R110)

IRT-SC5 (R055)

IRT-CS2 (R070)

IRT-CS22 (R047)

IRT-IPSF22 (R091)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

IRT-CS10 (R018)

IRT-IS10 (R110)

IRT-SC5 (R055)

IRT-CS2 (R070)

IRT-CS22 (R047)

IRT-IPSF22 (R091)

Clean: min Cp = -2.7 at x/c = 0.004

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

IRT-CS10 (R018)

IRT-IS10 (R110)

IRT-SC5 (R055)

IRT-CS2 (R070)

IRT-CS22 (R047)

IRT-IPSF22 (R091)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-CS10 (R018)

IRT-IS10 (R110)

IRT-SC5 (R055)

IRT-CS2 (R070)

IRT-CS22 (R047)

IRT-IPSF22 (R091)

3-22
(a) Lift coefficient (b) ΔCL vs α

(c) Drag coefficient (d) ΔCD vs α

(e) Pitching-moment coefficient (f) Hinge-moment coefficient 

FIGURE 3-38. EFFECT OF SMOOTH LEWICE ICE SHAPES ON CL, CD, CM, AND CH; 

LS-CS10, LS-IS10, LS-SC5, LS-CS2, LS-CS22N, LS-CS22S, AND LS-IPSF22 

CONFIGURATIONS; Re = 1.8x106; δA = 0°

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Lift coefficient

Clean (R006)

LS-CS10 (R334)

LS-IS10 (R172)

LS-SC5 (R187)

LS-CS2 (R206)

LS-CS22N (R414)

LS-CS22S (R442)

LS-IPSF22 (R348)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-90

-80

-70

-60

-50

-40

-30

-20

-10

0

10

20

ΔCL (%)

LS-CS10 (R334)

LS-IS10 (R172)

LS-SC5 (R187)

LS-CS2 (R206)

LS-CS22N (R414)

LS-CS22S (R442)

LS-IPSF22 (R348)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

0.00

0.05

0.10

0.15

0.20

0.25

0.30

0.35

0.40

Drag coefficient

Clean (R006)

LS-CS10 (R334)

LS-IS10 (R172)

LS-SC5 (R187)

LS-CS2 (R206)

LS-CS22N (R414)

LS-CS22S (R442)

LS-IPSF22 (R348)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-500

0

500

1000

1500

2000

2500

3000

ΔCD (%)

LS-CS10 (R334)

LS-IS10 (R172)

LS-SC5 (R187)

LS-CS2 (R206)

LS-CS22N (R414)

LS-CS22S (R442)

LS-IPSF22 (R348)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.10

-0.08

-0.06

-0.04

-0.02

0.00

0.02

0.04

0.06

0.08

0.10

Pitching moment coefficient

Clean (R006)

LS-CS10 (R334)

LS-IS10 (R172)

LS-SC5 (R187)

LS-CS2 (R206)

LS-CS22N (R414)

LS-CS22S (R442)

LS-IPSF22 (R348)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.12

-0.10

-0.08

-0.06

-0.04

-0.02

0.00

0.02

0.04

0.06

0.08

Hinge moment coefficient

Clean (R006)

LS-CS10 (R334)

LS-IS10 (R172)

LS-SC5 (R187)

LS-CS2 (R206)

LS-CS22N (R414)

LS-CS22S (R442)

LS-IPSF22 (R348)

3-23
(a) Cp vs x/c (α = -4°) (b) Cp vs x/c (α = 0°) 

(c) Cp vs x/c (α = 4°) (d) Cp vs x/c (α = 8°) 

(e) Cp vs x/c (α = 12°) (f) Cp vs x/c (α = 16°) 

FIGURE 3-39. EFFECT OF SMOOTH LEWICE ICE SHAPES ON PRESSURE 

DISTRIBUTIONS AT 15% SEMISPAN; LS-CS10, LS-IS10, LS-SC5, LS-CS2, LS-CS22N, 

LS-CS22S, AND LS-IPSF22 CONFIGURATIONS; RE = 1.8X106

; δA = 0°

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

LS-CS10 (R334)

LS-IS10 (R172)

LS-SC5 (R187)

LS-CS2 (R206)

LS-CS22N (R414)

LS-CS22S (R442)

LS-IPSF22 (R348)

Clean: max Cp = 0.7099 at x/c = 0

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

LS-CS10 (R334)

LS-IS10 (R172)

LS-SC5 (R187)

LS-CS2 (R206)

LS-CS22N (R414)

LS-CS22S (R442)

LS-IPSF22 (R348)

Clean: min Cp = -2.3254 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

LS-CS10 (R334)

LS-IS10 (R172)

LS-SC5 (R187)

LS-CS2 (R206)

LS-CS22N (R414)

LS-CS22S (R442)

LS-IPSF22 (R348)

Clean: min Cp = -3.2976 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

LS-CS10 (R334)

LS-IS10 (R172)

LS-SC5 (R187)

LS-CS2 (R206)

LS-CS22N (R414)

LS-CS22S (R442)

LS-IPSF22 (R348)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

-2.5

Pressure coefficient

Clean (R006)

LS-CS10 (R334)

LS-IS10 (R172)

LS-SC5 (R187)

LS-CS2 (R206)

LS-CS22N (R414)

LS-CS22S (R442)

LS-IPSF22 (R348)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

-2.5

-3.0

Pressure coefficient

Clean (R006)

LS-CS10 (R334)

LS-IS10 (R172)

LS-SC5 (R187)

LS-CS2 (R206)

LS-CS22N (R414)

LS-CS22S (R442)

LS-IPSF22 (R348)

3-24
(a) Cp vs x/c (α = -4°) (b) Cp vs x/c (α = 0°) 

(c) Cp vs x/c (α = 4°) (d) Cp vs x/c (α = 8°) 

(e) Cp vs x/c (α = 12°) (f) Cp vs x/c (α = 16°) 

FIGURE 3-40. EFFECT OF SMOOTH LEWICE ICE SHAPES ON PRESSURE 

DISTRIBUTIONS AT 50% SEMISPAN; LS-CS10, LS-IS10, LS-SC5, LS-CS2, LS-CS22N, 

LS-CS22S, AND LS-IPSF22 CONFIGURATIONS; Re = 1.8x106

; δA = 0°

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

LS-CS10 (R334)

LS-IS10 (R172)

LS-SC5 (R187)

LS-CS2 (R206)

LS-CS22N (R414)

LS-CS22S (R442)

LS-IPSF22 (R348)

Clean: min Cp = -3.1817 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

LS-CS10 (R334)

LS-IS10 (R172)

LS-SC5 (R187)

LS-CS2 (R206)

LS-CS22N (R414)

LS-CS22S (R442)

LS-IPSF22 (R348)

Clean: max Cp = 0.677 at x/c = 0.002

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

LS-CS10 (R334)

LS-IS10 (R172)

LS-SC5 (R187)

LS-CS2 (R206)

LS-CS22N (R414)

LS-CS22S (R442)

LS-IPSF22 (R348)

Clean: min Cp = -3.4743 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

LS-CS10 (R334)

LS-IS10 (R172)

LS-SC5 (R187)

LS-CS2 (R206)

LS-CS22N (R414)

LS-CS22S (R442)

LS-IPSF22 (R348)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

LS-CS10 (R334)

LS-IS10 (R172)

LS-SC5 (R187)

LS-CS2 (R206)

LS-CS22N (R414)

LS-CS22S (R442)

LS-IPSF22 (R348)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

LS-CS10 (R334)

LS-IS10 (R172)

LS-SC5 (R187)

LS-CS2 (R206)

LS-CS22N (R414)

LS-CS22S (R442)

LS-IPSF22 (R348)

3-25
(a) Cp vs x/c (α = -4°) (b) Cp vs x/c (α = 0°) 

(c) Cp vs x/c (α = 4°) (d) Cp vs x/c (α = 8°) 

(e) Cp vs x/c (α = 12°) (f) Cp vs x/c (α = 16°) 

FIGURE 3-41. EFFECT OF SMOOTH LEWICE ICE SHAPES ON PRESSURE 

DISTRIBUTIONS AT 85% SEMISPAN; LS-CS10, LS-IS10, LS-SC5, LS-CS2, LS-CS22N, 

LS-CS22S, AND LS-IPSF22 CONFIGURATIONS; Re = 1.8x106

; δA = 0°

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

LS-CS10 (R334)

LS-IS10 (R172)

LS-SC5 (R187)

LS-CS2 (R206)

LS-CS22N (R414)

LS-CS22S (R442)

LS-IPSF22 (R348)

Clean: max Cp = 0.6244 at x/c = 0.004

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

LS-CS10 (R334)

LS-IS10 (R172)

LS-SC5 (R187)

LS-CS2 (R206)

LS-CS22N (R414)

LS-CS22S (R442)

LS-IPSF22 (R348)

Clean: min Cp = -3.5329 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

LS-CS10 (R334)

LS-IS10 (R172)

LS-SC5 (R187)

LS-CS2 (R206)

LS-CS22N (R414)

LS-CS22S (R442)

LS-IPSF22 (R348)

Clean: min Cp = -2.7 at x/c = 0.004

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

LS-CS10 (R334)

LS-IS10 (R172)

LS-SC5 (R187)

LS-CS2 (R206)

LS-CS22N (R414)

LS-CS22S (R442)

LS-IPSF22 (R348)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

LS-CS10 (R334)

LS-IS10 (R172)

LS-SC5 (R187)

LS-CS2 (R206)

LS-CS22N (R414)

LS-CS22S (R442)

LS-IPSF22 (R348)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

LS-CS10 (R334)

LS-IS10 (R172)

LS-SC5 (R187)

LS-CS2 (R206)

LS-CS22N (R414)

LS-CS22S (R442)

LS-IPSF22 (R348)

3-26
(a) Lift coefficient (b) ΔCL vs α

(c) Drag coefficient (d) ΔCD vs α

(e) Pitching moment coefficient (f) Hinge-moment coefficient 

FIGURE 3-42. EFFECT OF ROUGH LEWICE ICE SHAPES ON CL, CD, CM, AND CH; 

LR-CS10, LR-IS10, LR-SC5, LR-CS2, LR-CS22N, LR-CS22S, AND LR-IPSF22 

CONFIGURATIONS; Re = 1.8x106

; δA = 0°

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

0.00

0.05

0.10

0.15

0.20

0.25

0.30

0.35

0.40

Drag coefficient

Clean (R006)

LR-CS10 (R363)

LR-IS10 (R240)

LR-SC5 (R216)

LR-CS2 (R232)

LR-CS22N (R459)

LR-CS22S (R468)

LR-IPSF22 (R374)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Lift coefficient

Clean (R006)

LR-CS10 (R363)

LR-IS10 (R240)

LR-SC5 (R216)

LR-CS2 (R232)

LR-CS22N (R459)

LR-CS22S (R468)

LR-IPSF22 (R374)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-100

-80

-60

-40

-20

0

20

40

60

80

ΔCL (%)

LR-CS10 (R363)

LR-IS10 (R240)

LR-SC5 (R216)

LR-CS2 (R232)

LR-CS22N (R459)

LR-CS22S (R468)

LR-IPSF22 (R374)

-10 -5 0 5 10 15 20 2

Angle of attack (deg)

5

-500

0

500

1000

1500

2000

2500

3000

3500

4000

ΔCD (%)

LR-CS10 (R363)

LR-IS10 (R240)

LR-SC5 (R216)

LR-CS2 (R232)

LR-CS22N (R459)

LR-CS22S (R468)

LR-IPSF22 (R374)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.12

-0.10

-0.08

-0.06

-0.04

-0.02

0.00

0.02

0.04

0.06

0.08

Hinge moment coefficient

Clean (R006)

LR-CS10 (R363)

LR-IS10 (R240)

LR-SC5 (R216)

LR-CS2 (R232)

LR-CS22N (R459)

LR-CS22S (R468)

LR-IPSF22 (R374)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.10

-0.08

-0.06

-0.04

-0.02

0.00

0.02

0.04

0.06

0.08

0.10

Pitching moment coefficient

Clean (R006)

LR-CS10 (R363)

LR-IS10 (R240)

LR-SC5 (R216)

LR-CS2 (R232)

LR-CS22N (R459)

LR-CS22S (R468)

LR-IPSF22 (R374)

3-27
0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

LR-CS10 (R363)

LR-IS10 (R240)

LR-SC5 (R216)

LR-CS2 (R232)

LR-CS22N (R459)

LR-CS22S (R468)

LR-IPSF22 (R374)

Clean: min Cp = -2.3254 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

LR-CS10 (R363)

LR-IS10 (R240)

LR-SC5 (R216)

LR-CS2 (R232)

LR-CS22N (R459)

LR-CS22S (R468)

LR-IPSF22 (R374)

Clean: max Cp = .7099 at x/c = 0

(a) Cp vs x/c (α = -4°) (b) Cp vs x/c (α = 0°) 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

LR-CS10 (R363)

LR-IS10 (R240)

LR-SC5 (R216)

LR-CS2 (R232)

LR-CS22N (R459)

LR-CS22S (R468)

LR-IPSF22 (R374)

Clean: min Cp = -3.2976 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

LR-CS10 (R363)

LR-IS10 (R240)

LR-SC5 (R216)

LR-CS2 (R232)

LR-CS22N (R459)

LR-CS22S (R468)

LR-IPSF22 (R374)

(c) Cp vs x/c (α = 4°) (d) Cp vs x/c (α = 8°) 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

-2.5

-3.0

Pressure coefficient

Clean (R006)

LR-CS10 (R363)

LR-IS10 (R240)

LR-SC5 (R216)

LR-CS2 (R232)

LR-CS22N (R459)

LR-CS22S (R468)

LR-IPSF22 (R374)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

LR-CS10 (R363)

LR-IS10 (R240)

LR-SC5 (R216)

LR-CS2 (R232)

LR-CS22N (R459)

LR-CS22S (R468)

LR-IPSF22 (R374)

(e) Cp vs x/c (α = 12°) (f) Cp vs x/c (α = 16°) 

FIGURE 3-43. EFFECT OF ROUGH LEWICE ICE SHAPES ON PRESSURE 

DISTRIBUTIONS AT 15% SEMISPAN; LR-CS10, LR-IS10, LR-SC5, LR-CS2, LR-CS22N, 

LR-CS22S, AND LR-IPSF22 CONFIGURATIONS; Re = 1.8x106

; δA = 0°

3-28
0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

LR-CS10 (R363)

LR-IS10 (R240)

LR-SC5 (R216)

LR-CS2 (R232)

LR-CS22N (R459)

LR-CS22S (R468)

LR-IPSF22 (R374)

Clean: max Cp = 0.677 at x/c = 0.002

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

LR-CS10 (R363)

LR-IS10 (R240)

LR-SC5 (R216)

LR-CS2 (R232)

LR-CS22N (R459)

LR-CS22S (R468)

LR-IPSF22 (R374)

Clean: min Cp = -3.1817 at x/c = 0.005

(a) Cp vs x/c (α = -4°) (b) Cp vs x/c (α = 0°) 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

LR-CS10 (R363)

LR-IS10 (R240)

LR-SC5 (R216)

LR-CS2 (R232)

LR-CS22N (R459)

LR-CS22S (R468)

LR-IPSF22 (R374)

Clean: min Cp = -3.4743 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

LR-CS10 (R363)

LR-IS10 (R240)

LR-SC5 (R216)

LR-CS2 (R232)

LR-CS22N (R459)

LR-CS22S (R468)

LR-IPSF22 (R374)

(c) Cp vs x/c (α = 4°) (d) Cp vs x/c (α = 8°) 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

LR-CS10 (R363)

LR-IS10 (R240)

LR-SC5 (R216)

LR-CS2 (R232)

LR-CS22N (R459)

LR-CS22S (R468)

LR-IPSF22 (R374)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

LR-CS10 (R363)

LR-IS10 (R240)

LR-SC5 (R216)

LR-CS2 (R232)

LR-CS22N (R459)

LR-CS22S (R468)

LR-IPSF22 (R374)

(e) Cp vs x/c (α = 12°) (f) Cp vs x/c (α = 16°) 

FIGURE 3-44. EFFECT OF ROUGH LEWICE ICE SHAPES ON PRESSURE 

DISTRIBUTIONS AT 50% SEMISPAN; LR-CS10, LR-IS10, LR-SC5, LR-CS2, LR-CS22N, 

LR-CS22S, AND LR-IPSF22 CONFIGURATIONS; Re = 1.8x106

; δA = 0°

3-29
0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

LR-CS10 (R363)

LR-IS10 (R240)

LR-SC5 (R216)

LR-CS2 (R232)

LR-CS22N (R459)

LR-CS22S (R468)

LR-IPSF22 (R374)

Clean: min Cp = -3.5329 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

LR-CS10 (R363)

LR-IS10 (R240)

LR-SC5 (R216)

LR-CS2 (R232)

LR-CS22N (R459)

LR-CS22S (R468)

LR-IPSF22 (R374)

Clean: max Cp = 0.6244 at x/c = 0.004

(a) Cp vs x/c (α = -4°) (b) Cp vs x/c (α = 0°) 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

LR-CS10 (R363)

LR-IS10 (R240)

LR-SC5 (R216)

LR-CS2 (R232)

LR-CS22N (R459)

LR-CS22S (R468)

LR-IPSF22 (R374)

Clean: min Cp = -2.7 at x/c = 0.004

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

LR-CS10 (R363)

LR-IS10 (R240)

LR-SC5 (R216)

LR-CS2 (R232)

LR-CS22N (R459)

LR-CS22S (R468)

LR-IPSF22 (R374)

(c) Cp vs x/c (α = 4°) (d) Cp vs x/c (α = 8°) 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

LR-CS10 (R363)

LR-IS10 (R240)

LR-SC5 (R216)

LR-CS2 (R232)

LR-CS22N (R459)

LR-CS22S (R468)

LR-IPSF22 (R374)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

LR-CS10 (R363)

LR-IS10 (R240)

LR-SC5 (R216)

LR-CS2 (R232)

LR-CS22N (R459)

LR-CS22S (R468)

LR-IPSF22 (R374)

(e) Cp vs x/c (α = 12°) (f) Cp vs x/c (α = 16°) 

FIGURE 3-45. EFFECT OF ROUGH LEWICE ICE SHAPES ON PRESSURE 

DISTRIBUTIONS AT 85% SEMISPAN; LR-CS10, LR-IS10, LR-SC5, LR-CS2, LR-CS22N, 

LR-CS22S, AND LR-IPSF22 CONFIGURATIONS; Re = 1.8x106

; δA = 0°

3-30
TABLE 3-1. EFFECT OF IRT ICE SHAPES ON CL AND CD; ReMAC = 1.8x106

; δA = 0°

α (deg) Configuration CL ΔCL CD ΔCD

CLEAN 0.22 0.0% 0.008 0.0% 

IRT-CS10 0.19 -13.6% 0.074 825.0% 

IRT-IS10 0.20 -9.1% 0.050 525.0% 

IRT-SC5 0.22 0.0% 0.016 100.0% 

IRT-CS2 0.21 -4.5% 0.021 162.5% 

IRT-CS22 0.05 -77.3% 0.224 2700.0% 

3 

IRT-IPSF22 0.18 -18.2% 0.081 912.5% 

CLEAN 0.38 0.0% 0.014 0.0% 

IRT-CS10 0.35 -7.9% 0.089 535.7% 

IRT-IS10 0.37 -2.6% 0.064 357.1% 

IRT-SC5 0.38 0.0% 0.023 64.3% 

IRT-CS2 0.37 -2.6% 0.029 107.1% 

IRT-CS22 0.05 -86.8% 0.229 1535.7% 

5 

IRT-IPSF22 0.34 -10.5% 0.097 592.9% 

CLEAN 0.76 0.0% 0.053 0.0% 

IRT-CS10 0.54 -28.9% 0.165 211.3% 

IRT-IS10 0.64 -15.8% 0.154 190.6% 

IRT-SC5 0.74 -2.6% 0.084 58.5% 

IRT-CS2 0.72 -5.3% 0.100 88.7% 

IRT-CS22 0.21 -72.4% 0.261 392.5% 

10 

IRT-IPSF22 0.53 -30.3% 0.172 224.5% 

CLEAN 0.86 0.0% 0.198 0.0% 

IRT-CS10 0.60 -30.2% 0.238 20.2% 

IRT-IS10 0.61 -29.1% 0.224 13.1% 

IRT-SC5 0.90 4.7% 0.221 11.6% 

IRT-CS2 0.76 -11.6% 0.209 5.6% 

IRT-CS22 0.41 -52.3% 0.323 63.1% 

15 

IRT-IPSF22 0.56 -34.9% 0.234 18.3% 

TABLE 3-2. EFFECT OF IRT ICE SHAPES ON CL,stall, αstall, AND 

CD,min; ReMAC = 1.8x106

; δA = 0°

Configuration CL,stall ΔCL,stall αstall Δαstall

CL at 

α = 13.8°

ΔCL at

α = 13.8° CD,min ΔCD,min

CD at 

α = 13.8°

ΔCD at

α = 13.8°

Clean 0.87 0.0% 13.8° 0.0° 0.87 0.0% 0.006 0.0% 0.147 0.0% 

IRT-CS10 0.54 -37.9% 10.5° -3.3° 0.56 -35.6% 0.072 1100.0% 0.209 42.2% 

IRT-IS10 0.64 -26.4% 10.6° -3.2° 0.59 -32.2% 0.047 683.3% 0.198 34.7% 

IRT-SC5 0.90 3.4% 15.8° 2.0° 0.86 -1.1% 0.014 133.3% 0.172 17.0% 

IRT-CS2 0.77 -11.5% 12.7° -1.1° 0.76 -12.6% 0.018 200.0% 0.174 18.4% 

IRT-CS22 0.06 -93.6% 6.0° -7.8° 0.36 -58.6% 0.218 3533.3% 0.300 104.1% 

IRT-IPSF22 0.53 -39.1% 10.5° -3.3° 0.53 -39.1% 0.078 1200.0% 0.210 42.9% 

3-31
TABLE 3-3. EFFECT OF IRT ICE SHAPES ON HINGE-MOMENT COEFFICIENT SLOPE; 

ReMAC = 1.8x106

; δA = 0°

Region A Region B Region C 

Ice shape 

dCH/dα

(per deg) 

α-range 

(deg) 

dCH/dα

(per deg) 

α-range 

(deg) 

dCH/dα

(per deg) 

α-range 

(deg) 

Clean -0.0026 0 to 10 -0.0155 10 to 14 -0.0001 14 to 20 

IRT-CS10 -0.0038 0 to 6 -0.0091 6 to 10 -0.0019 10 to 20 

IRT-IS10 -0.0032 0 to 6 -0.0106 6 to 10 -0.0016 10 to 20 

IRT-SC5 -0.0022 0 to 9 -0.0143 9 to 15 0.0004 15 to 20 

IRT-CS2 -0.0023 0 to 9 -0.0137 9 to 13 -0.0016 13 to 20 

IRT-CS22 -0.0057 0 to 6 -0.0065 6 to 11 -0.0021 11 to 20 

IRT-IPSF22 -0.0041 0 to 6 -0.0100 6 to 10 -0.0017 10 to 20 

TABLE 3-4. EFFECT OF SMOOTH LEWICE ICE SHAPES ON CL AND CD; 

ReMAC = 1.8x106

; δA = 0°

α (deg) Configuration CL ΔCL CD ΔCD

CLEAN 0.22 0.0% 0.008 0.0% 

LS-CS10 0.18 -18.2% 0.079 887.5% 

LS-IS10 0.22 0.0% 0.040 400.0% 

LS-SC5 0.22 0.0% 0.012 50.0% 

LS-CS2 0.22 0.0% 0.020 150.0% 

LS-CS22N 0.16 -27.3% 0.135 1587.5% 

LS-CS22S 0.12 -45.5% 0.155 1837.5% 

3 

LS-IPSF22 0.22 0.0% 0.071 787.5% 

CLEAN 0.38 0.0% 0.014 0.0% 

LS-CS10 0.34 -10.5% 0.093 564.3% 

LS-IS10 0.39 2.6% 0.058 314.3% 

LS-SC5 0.39 2.6% 0.017 21.4% 

LS-CS2 0.39 2.6% 0.030 114.3% 

LS-CS22N 0.25 -34.2% 0.156 1014.3% 

LS-CS22S 0.21 -44.7% 0.168 1100.0% 

5 

LS-IPSF22 0.36 -5.3% 0.088 528.6% 

CLEAN 0.76 0.0% 0.053 0.0% 

LS-CS10 0.61 -19.7% 0.171 222.6% 

LS-IS10 0.63 -17.1% 0.148 179.2% 

LS-SC5 0.79 3.9% 0.048 -9.4% 

LS-CS2 0.69 -9.2% 0.120 126.4% 

LS-CS22N 0.43 -43.4% 0.233 339.6% 

LS-CS22S 0.43 -43.4% 0.238 349.1% 

10 

LS-IPSF22 0.61 -19.7% 0.169 218.9% 

CLEAN 0.86 0.0% 0.198 0.0% 

LS-CS10 0.71 -17.4% 0.264 33.3% 

LS-IS10 0.67 -22.1% 0.231 16.7% 

LS-SC5 0.93 8.1% 0.189 -4.5% 

LS-CS2 0.74 -14.0% 0.217 9.6% 

LS-CS22N 0.50 -41.9% 0.313 58.1% 

LS-CS22S 0.55 -36.0% 0.319 61.1% 

15 

LS-IPSF22 0.68 -20.9% 0.256 29.3% 

3-32
TABLE 3-5. EFFECT OF SMOOTH LEWICE ICE SHAPES ON CL,stall, αstall, AND CDmin; 

ReMAC =1.8x106

; δA = 0°

Configuration CL,stall ΔCL,stall αstall Δαstall

CL at 

α=13.8°

ΔCL at 

α=13.8° CD,min ΔCD,min

CD at 

α=13.8°

ΔCD at 

α=13.8°

Clean 0.87 0.0% 13.8° 0.0% 0.87 0.0% 0.006 0.0% 0.147 0.0% 

LS-CS10 0.71 -18.4% 14.6° 5.8% 0.70 -19.5% 0.078 1200.0% 0.229 55.8% 

LS-IS10 0.68 -21.8% 13.6° -1.4% 0.68 -21.8% 0.036 500.0% 0.202 37.4% 

LS-SC5 0.94 8.0% 14.8° 7.2% 0.93 6.9% 0.010 66.7% 0.124 -15.6% 

LS-CS2 0.75 -13.8% 14.7° 6.5% 0.74 -14.9% 0.017 183.3% 0.182 23.8% 

LS-CS22N 0.49 -43.7% 13.5° -2.2% 0.49 -43.7% 0.128 2033.3% 0.289 96.6% 

LS-CS22S 0.54 -37.9% 13.4° -2.9% 0.54 -37.9% 0.148 2366.7% 0.294 100.0% 

LS-IPSF22 0.69 -20.7% 14.6° 5.8% 0.68 -21.8% 0.066 1000.0% 0.225 53.1% 

TABLE 3-6. EFFECT OF SMOOTH LEWICE ICE SHAPES ON HINGE-MOMENT 

COEFFICIENT SLOPE; ReMAC = 1.8x106

; δA = 0°

Region A Region B Region C 

Ice Shape 

dCH/dα

(per deg) 

α-range 

(deg) 

dCH/dα

(per deg) 

α-range 

(deg) 

dCH/dα

(per deg) 

α-range 

(deg) 

Clean -0.0026 0 to 10 -0.0155 10 to 14 -0.0001 14 to 20 

LS-CS10 -0.0134 0 to 7 -0.0014 7 to 15 -0.0024 14 to 20 

LS-IS10 -0.0067 0 to 7 -0.0033 7 to 13 -0.0023 13 to 20 

LS-SC5 -0.0027 0 to 11 -0.0175 11 to 14 0.0003 14 to 20 

LS-CS2 -0.0031 0 to 8 -0.0068 8 to 14 -0.0017 14 to 20 

LS-CS22N -0.0089 0 to 6 -0.0039 6 to 14 -0.0060 14 to 20 

LS-CS22S -0.0095 0 to 7 -0.0045 7 to 13 -0.0016 13 to 20 

LS-IPSF22 -0.0109 0 to 7 -0.0014 7 to 14 -0.0021 14 to 20 

3-33
TABLE 3-7. EFFECT OF ROUGH LEWICE ICE SHAPES ON CL AND CD; 

ReMAC = 1.8x106

; δA = 0°

α (deg) Configuration CL ΔCL CD ΔCD

CLEAN 0.22 0.0% 0.008 0.0% 

LR-CS10 0.18 -18.2% 0.077 862.5% 

LR-IS10 0.22 0.0% 0.043 437.5% 

LR-SC5 0.21 -4.5% 0.014 75.0% 

LR-CS2 0.21 -4.5% 0.023 187.5% 

LR-CS22N 0.17 -22.7% 0.189 2262.5% 

LR-CS22S 0.15 -31.8% 0.214 2575.0% 

3 

LR-IPSF22 0.21 -4.5% 0.070 775.0% 

CLEAN 0.38 0.0% 0.014 0.0% 

LR-CS10 0.33 -13.2% 0.091 550.0% 

LR-IS10 0.38 0.0% 0.060 328.6% 

LR-SC5 0.38 0.0% 0.019 35.7% 

LR-CS2 0.38 0.0% 0.032 128.6% 

LR-CS22N 0.24 -36.8% 0.202 1342.9% 

LR-CS22S 0.21 -44.7% 0.225 1507.1% 

5 

LR-IPSF22 0.36 -5.3% 0.089 535.7% 

CLEAN 0.76 0.0% 0.053 0.0% 

LR-CS10 0.55 -27.6% 0.168 217.0% 

LR-IS10 0.61 -19.7% 0.147 177.4% 

LR-SC5 0.76 0.0% 0.058 9.4% 

LR-CS2 0.69 -9.2% 0.121 128.3% 

LR-CS22N 0.40 -47.4% 0.254 379.2% 

LR-CS22S 0.34 -55.3% 0.266 401.9% 

10 

LR-IPSF22 0.57 -25.0% 0.167 215.1% 

CLEAN 0.86 0.0% 0.198 0.0% 

LR-CS10 0.61 -29.1% 0.249 25.8% 

LR-IS10 0.63 -26.7% 0.226 14.1% 

LR-SC5 0.98 14.0% 0.212 7.1% 

LR-CS2 0.72 -16.3% 0.218 10.1% 

LR-CS22N 0.41 -52.3% 0.307 55.1% 

LR-CS22S 0.35 -59.3% 0.311 57.1% 

15 

LR-IPSF22 0.61 -29.1% 0.246 24.2% 

TABLE 3-8. EFFECT OF ROUGH LEWICE ICE SHAPES ON CL,stall, 

αstall, AND CD,min; ReMAC = 1.8x106

, δA = 0°

Configuration CL,stall ΔCL,stall αstall Δαstall

CL at 

α = 13.8°

ΔCL at 

α = 13.8° CD,min ΔCD,min

CD at 

α = 13.8°

ΔCD at

α = 13.8°

Clean 0.87 0.0% 13.8° 0.0° 0.87 0.0% 0.006 0.0% 0.147 0.0% 

LR-CS10 0.61 -29.9% 13.6° -0.2° 0.61 -29.9% 0.076 1166.7% 0.221 50.3% 

LR-IS10 0.65 -25.3% 13.6° -0.2° 0.64 -26.4% 0.039 550.0% 0.200 36.1% 

LR-SC5 0.98 12.6% 15.8° 2.0° 0.92 5.7% 0.013 116.7% 0.146 -0.7% 

LR-CS2 0.72 -17.2% 13.7° -0.1° 0.72 -17.2% 0.020 233.3% 0.184 25.2% 

LR-CS22N 0.41 -52.9% 11.4° -2.4° 0.40 -54.0% 0.182 2933.3% 0.291 98.0% 

LR-CS22S 0.34 -60.9% 10.3° -3.5° 0.33 -62.1% 0.205 3316.7% 0.295 100.7% 

LR-IPSF22 0.61 -29.9% 12.6° -1.2° 0.61 -29.9% 0.065 983.3% 0.218 48.3% 

3-34
TABLE 3-9. EFFECT OF ROUGH LEWICE ICE SHAPES ON HINGE-MOMENT 

COEFFICIENT SLOPE; ReMAC = 1.8x106

, δA = 0°

Region A Region B Region C 

Ice Shape 

dCH/dα

(per deg) 

α-range 

(deg) 

dCH/dα

(per deg) 

α-range 

(deg) 

dCH/dα

(per deg) 

α-range 

(deg) 

Clean -0.0026 0 to 10 -0.0155 10 to 14 -0.0001 14 to 20 

LR-CS10 -0.0164 0 to 6 -0.0017 6 to 13 -0.0023 13 to 20 

LR-IS10 -0.0062 0 to 6 -0.0040 6 to 13 -0.0021 13 to 20 

LR-SC5 -0.0026 0 to 11 -0.0157 11 to 15 0.0005 15 to 20 

LR-CS2 -0.0032 0 to 8 -0.0101 8 to 12 -0.0016 12 to 20 

LR-CS22N 0.0007 0 to 7 -0.0036 7 to 11 -0.0100 11 to 20 

LR-CS22S 0.0007 0 to 5 0.0009 5 to 10 -0.0120 10 to 20 

LR-IPSF22 -0.0160 0 to 6 -0.0013 6 to 12 -0.0025 12 to 20 

3.2.3.1 Clean Wing. 

Maximum lift coefficient for the clean wing was 0.87 and occurred at an α of 13.8°, as shown in 

figure 3-34(a) and in table 3-2. The lift dropped gradually after stall and was reduced to about 

0.81 at α of 20°. The slope of the linear portion of the lift curve was 4.41 per radian or 0.077 per 

degree. This slope correlates well with the slope of 0.064 obtained from equation 3-5 [16]. 

( )2

2

0

2

21

0

0

cos ∞ ⎟ ⋅− ⎠

⎞ ⎜

⎝

⎛ + ⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

Λ ⎟ +

⎠

⎞ ⎜

⎝

⎛

⋅ =

MAR a AR a

ARa CL

π π

α

(3-5)

where a0 is the 2D lift slope of the GLC-305 airfoil and is equal to 0.084 per degree based on the 

data of reference 17, Λ1/2 is the wing sweep angle at 50% chord and is equal to 22°, AR is the 

wing aspect ratio (6.8) given in figure 2-3 and M∞ is the free-stream Mach number, which for 

the WSU wind tunnel tests was 0.185, as shown in table 2-4. 

From examination of the pressure distributions and the flow visualization data obtained, flow 

separation was initiated at the wing LE and was combined with a leading-edge vortex. Flow 

separation was first observed for α of 9° near the 30% semispan station. The region of flow 

separation was very small in both the spanwise and chordwise directions (2% chord and about 

3% semispan). As α was increased to 10°, a large trapezoidal region of rough and separated 

flow was observed between the 30% and 85% semispan stations and between 12% and 60% of 

wing chord. At α of 11°, considerable TE separation occurred near the wing trailing edge over a 

chord length of about 20%. Trailing edge separation extended from about 55% semispan to the 

wing tip. At α of 12°, flow separation was observed over the wing tip (outboard 15% of 

semispan) and near the TE of the inboard (5% to 30% semispan) portion of the wing. Finally, at 

α between 13° and 14°, complete flow separation occurred over most of the wing upper surface 

which led to wing stall. 

Drag performance for the clean wing is presented in figure 3-34(c) and in tables 3-1 and 3-2. 

The minimum drag coefficient was 0.006 at α of 1°. The drag increased to a maximum value at 

3-35
about 0.28 at α of 20°. The maximum lift to drag ratio for the swept wing was 27.5 at α of 3.2°. 

Pitching moment about the 25% MAC location was nearly flat, as shown in figure 3-34(e), until 

about stall where a considerable increase in negative pitching moment (leading edge down) was 

observed as the load center moved downstream of the quarter-chord point due to flow separation. 

Pitching-moment coefficient ranged from -0.187 to 0.15. The pitching moment is slightly stable 

and there is a tendency for pitch up at α of 10° to 12°. The pitching moment then breaks with 

positive stability, airplane nose down, following stall and α in the range 15° to 17°. 

Hinge-moment coefficients are presented in figure 3-34(f) and in table 3-3. The maximum slope 

of the hinge-moment coefficient occurred in region B and was -0.0155 per degree. This 

corresponds to a change in hinge moment of 0.52 in-lb per degree for the wing model tested. For 

a control deflection of 0°, the hinge-moment coefficient was negative (i.e., leading edge down) 

for positive C as expected. The graph of hinge moment coefficient versus α is linear in the 

region corresponding to the linear part of the lift curve and breaks as maximum lift is approached 

and the hinge moment slightly reverses at α larger than 15°. 

3.2.3.2 Icing Research Tunnel Ice Shape Castings. 

3.2.3.2.1 Lift Coefficient. 

Ice shapes caused significant changes in the wing lift characteristics, as shown in figures 3-34(a) 

and 3-34(b). With the exception of the IRT-SC5 ice shape, which increased CL,stall and αstall with 

respect to the clean wing, all ice shapes reduced lift throughout the α-range. In addition, the iced 

wing αstall and the linear lift slope were reduced with respect to the clean configuration. 

Table 3-1 provides lift performance for the clean and iced wing cases for angles of attack 

corresponding to the linear and nonlinear portions of the lift curve. Table 3-2 compares CL,stall

and αstall for all six ice shapes tested. The CL,stall reduction was in the range of 11.5% to 93.6% 

and the corresponding reduction in αstall ranged from 8% to 56%. The largest degradation in lift 

performance was obtained with the IRT-CS22 ice shape. For this ice shape, the large upper and 

lower horns near the wing leading edge caused extensive flow separation, even at low α. The 

improved lift performance observed with the IRT-SC5 ice shape was mainly due to the small 

leading-edge droop (LE flap effect) caused by the ice shape. The pressure distributions at the 

50% semispan station presented in figure 3-36 demonstrate that for α of 8°, the suction peak for 

the IRT-SC5 case was lower than the clean wing, indicating a lower effective α. In addition, 

figure 3-36(f) shows that at α of 16°, the baseline experienced extensive flow separation, while 

flow separation for the wing with the IRT-SC5 ice shape was considerably reduced. 

The aerodynamic performance data for the IRT-CS2, IRT-CS10, and IRT-CS22 complete 

scallop glaze ice shapes indicate that the ice shapes with the larger horn size resulted in greater 

penalties in lift characteristics, as shown in figures 3-34(a) and 3-34(b). The iced wing 

performance in terms of αstall and CL,stall was progressively reduced in the following ice shape 

sequence: IRT-CS2, IRT-CS10, and IRT-CS22. 

3-36
3.2.3.2.2 Drag Coefficient. 

The increase in the drag coefficient due to the ice shapes tested is demonstrated in 

figures 3-34(c) and 3-34(d) and in tables 3-1 and 3-2. In general, the increase in minimum drag 

due to the ice accretions ranged from 1.3 to 35 times that of the clean wing. The smallest drag 

rise was caused by the IRT-SC5 ice shape, while the largest drag increment was due to the 

IRT-CS22 wing. Near the αstall of the clean wing, the drag due to the ice shapes was 17% to 

104% greater than the clean wing. Furthermore, as the ice shape horn size increased (from 

IRT-CS2 to IRT-CS22 ice shapes), the drag increment of the iced wing increased, due to 

increased flow separation downstream of the ice shape. 

3.2.3.2.3 Pitching-Moment Coefficient. 

All pitching moment data presented in figure 3-34(e) are about the 25% MAC point. The 

addition of the ice shapes caused considerable changes in CM,MAC/4 due to the shift in the load 

distribution caused by the separated flow downstream of the ice shapes. In general, for positive 

α, the ice shapes caused more positive pitching moment prior to stall compared to the clean wing 

case. Thus, the lift vector for the iced wing was upstream of the 25% MAC location. The ice 

shape causes unstable behavior, with a break at α of 6°. Stability increases for α greater than 6°. 

Note that a 0.01 change in the value of the pitching-moment coefficient corresponds to a change 

of 5.7 ft-lb in pitching moment about the MAC/4 point. 

3.2.3.2.4 Hinge-Moment Coefficient. 

The differences observed between the clean and iced wing cases in figure 3-34(f) were mainly 

due to the increased separation over the aileron upper surface caused by the ice shapes. In 

general, the ice shapes moved the start of region B to the left (lower α) and increased the hingemoment coefficients over region B (CH became more negative). In all cases, the maximum hinge 

moment for the iced wing in region C was bounded by the maximum CH of the clean wing, as 

shown in figure 3-34(f). Note a shift in the entire CH versus α plot. A control force reversal 

(i.e., change in CH from positive to negative) was observed between α of 0° and 3° for the 10- 

and 22.5-min ice shapes. This was caused by increased flow separation over the lower surface of 

the wing that resulted in greater suction over the control lower surface. Thus, the aileron had the 

tendency to move trailing edge down for α between 0° and 3°. 

The effect of horn height on aileron hinge moments can be assessed by reviewing the results 

obtained with the 2-min (IRT-CS2), 10-min (IRT-CS10), and 22.5-min (IRT-CS22) complete 

scallop glaze ice shapes. As demonstrated in table 3-3, the slope of the CH curve in region B 

decreased as the ice horn height was decreased. However, in regions A and C, the CH slope 

increased as the ice shape horn height was increased. 

3.2.3.2.5 Pressure Distributions. 

Pressure distributions corresponding to wing sections at the 15%, 50%, and 85% semispan 

locations are presented in figures 3-35, 3-36, and 3-37, respectively. The pressures are for α of 

-4°, 0°, 4°, 8°, 12°, and 16°, which cover the linear and nonlinear lift range. The results indicate 

3-37
that, in most cases, the addition of ice shapes resulted in a dramatic change in the clean wing 

pressure distribution. Near the wing LE, surface pressures for the iced wing featured a region of 

separated flow, which was characterized by a flat pressure distribution followed by pressure 

recovery. The region under the flat curve in the pressure data indicates the presence of a 

separated flow bubble. The extent of bubble was a function of ice shape and α. In many cases, 

particularly near the 50% and 85% spanwise stations, massive flow separation was observed as α

was increased. Massive flow separation was associated with flat pressure distributions that 

extended to the wing TE. At high positive α, separated flow was observed over the wing upper 

surface. For ice shape with the large horns, extensive flow separation occurred over both wing 

surfaces at low α. Flow separation caused considerable changes in the load distribution over the 

wing and the control surface, which was the main reason for the observed changes in 

aerodynamic performance and aileron hinge moment with respect to the clean wing. Pressure 

trends for the iced configurations and for each α presented in figures 3-35, 3-36, and 3-37 are 

summarized below: 

• α = -4°: In general, suction was observed over the lower surface of the wing and was 

associated with a leading-edge bubble. For all ice shapes, the bubble extent increased 

towards the outboard sections of the wing. In most cases, the highest suction occurred 

near the inboard sections of the wing (i.e., 0% to 30% of semispan). The IRT-CS22 ice 

shape caused massive flow separation at practically all spanwise stations. 

• α = 0°: Suction remained higher over the lower surface for practically all ice shapes. 

Once again, LE bubbles were evident at all three spanwise locations for most of the ice 

shapes tested. The bubble extent was a function of horn size. Large horns resulted in 

longer bubbles. For the IRT-CS10, IRT-IS10, IRT-CS22, and IRT-IPSF22 ice shapes, 

the maximum suction occurred at spanwise stations located between 15% and 50% 

semispan. For the small ice accretions, namely IRT-SC5 and IRT-CS2, the maximum 

suction took place over the outboard part of the wing. Note that at α of 0°, the wing tip 

was at a geometric α of -4o due to the wing twist. 

• α = 4°: For this angle of attack, suction for most of the ice shape cases was increased 

over the wing upper surface. The Cp on the lower surface varied from negative near the 

LE, indicating the presence of separation bubbles, to low negative and, in some cases, 

positive near the TE. The extent of the upper surface bubbles varied with spanwise 

distance from root to tip due to the negative wing twist, which resulted in a lower αlocal

over the outboard sections. For the 10- and 22.5-min ice shapes, the extent of the 

separation bubbles was, in most cases, greater over the middle and outboard sections of 

the wing. For these ice shapes, maximum suction did not vary significantly with 

spanwise location. 

• α = 8°: Upper surface suction and bubble extent was increased at this α at all spanwise 

stations for ice shapes IRT-CS10, IRT-IS10, IRT-CS2, IRT-CS22, and IRT-IPSF22. For 

the 10- and 22.5-min ice shapes, extensive flow separation was observed at the 85% 

semispan station. The 5-min rime ice shape, IRT-SC5, did not exhibit bubble formation 

over the two inboard stations. However, at the outboard station (85% semispan), a small 

bubble extending to about 20% chord was observed. Maximum upper surface suction 

3-38
took place over the wing near 15% semispan. The only exception was the IRT-CS22 

case for which the maximum suction occurred at the 85% semispan station. With the 

exception of the IRT-CS22 ice shape, all other ice shapes had positive or low negative 

pressure coefficients over the lower surface of the wing. 

• α = 12°: Large bubbles were observed at the 15% semispan station and complete flow 

separation was evident over the wing upper surface at the 50% and 85% semispan 

locations with the 10- and 22.5-min ice shapes tested. The flow over the IRT-SC5 and 

IRT-CS2 ice shapes exhibited smaller bubbles compared to the larger ice shapes. The 

bubbles for these two ice accretions extended from about 20% to about 75% chord 

length, depending on ice shape and spanwise station. For all ice shapes, maximum 

suction was observed at the inboard station located at 15% semispan. 

• α = 16°: At this angle of attack, massive flow separation was evident over the wing 

upper surface at all spanwise stations for all 10- and 22.5-min ice shapes. In all cases, 

maximum suction took place at the 15% semispan station. The 2- and 5-min ice 

accretions exhibited long bubbles over the inboard stations. However, at the outboard 

stations, the flow for these two ice shapes was completely separated over the wing upper 

surface. 

The impact of the horn size on aerodynamic performance can be explained by reviewing the 

pressure distributions. Specifically for ice shapes with large horn heights, LE separation bubbles 

were observed at low α. These bubbles became progressively longer in the chordwise direction 

and eventually burst to form a region of massive flow separation. Ice shapes of smaller horn size 

formed bubbles with smaller chordwise extent. These bubbles did not appear until α was higher 

with respect to where the bubbles were observed with the larger horn ice shapes. With the IRTCS22 configuration, extensive flow separation occurred over the wing suction surface even at 

low α. Although the flow over the upper surface of wing with the IRT-CS22 ice shape remained 

separated throughout the positive α-range, the lift increased as α was increased beyond about 6o. 

This was due to the increase in the pressure on the lower surface of the wing where the flow 

remained mostly attached. Pressure on the upper surface did not vary significantly with α once 

the wing had stalled. 

3.2.3.3 Smooth LEWICE Ice Shapes. 

The effects of the smooth LEWICE ice shapes on wing performance are demonstrated in 

figures 3-38 to 3-41 and in tables 3-4 to 3-6. 

3.2.3.3.1 Lift Coefficient. 

With the exception of the LS-SC5 ice shape, which increased CL,stall and αstall, all ice shapes 

reduced lift performance with respect to the clean wing, as demonstrated in figures 3-38(a) and 

3-38(b) and in tables 3-4 and 3-5. In assessing these reductions note that they could be larger if 

clean wing data were available for higher Reynolds numbers. The iced wing lift curves exhibited 

reduced lift coefficient and lower lift slope with respect to that of the clean wing. The reduction 

of CL,stall ranged from 18.4% to 43.7%. The largest loss in lift was obtained with the LS-CS22N 

ice shape. As shown in table 3-5, smooth LEWICE ice shapes LS-CS10, LS-SC5, LS-CS2, and 

3-39
LS-IPSF22 increased αstall by 5.8% to 7.2%, while LS-IS10, LS-CS22N, and LS-CS22S ice 

shapes reduced αstall by 1.4% to 2.9%. The 5-min rime ice shape LS-SC5 increased CL,stall by 7% 

and increased αstall by 1°. The observed performance increment in the iced wing lift was due to 

the leading-edge droop (i.e., leading-edge flap effect) caused by the 5-min rime ice shape. 

The effect of horn height on the lift performance can be observed by comparing the iced wing lift 

performance with the LS-CS2, LS-CS10, LS-CS22N, and LS-CS22S ice shapes. The CL,stall

reduction with respect to the clean wing was -13.8%, -18.4%, -43.7%, and -38% for the LS-CS2, 

LS-CS10, LS-CS22N, and LS-CS22S ice shapes respectively. The corresponding change in αstall 

was 6.5%, 5.8%, -2.2%, and -2.9% for the same sequence of ice shapes. 

3.2.3.3.2 Drag Coefficient. 

In general, all smooth LEWICE ice shapes increased wing drag considerably, as demonstrated in 

figures 3-38(c) and 3-38(d) and in tables 3-4 and 3-5. The results presented show that the 

increase in CD,min due to the ice shapes ranged from 66.7% for ice shape LS-SC5 to 2366.7% for 

ice shape LS-CS22S with respect to the clean wing. For α near the clean wing αstall, drag rise for 

the iced wing was in the range of -15.6% for the LS-SC5 to 100% for LS-CS22S with respect to 

the clean case. The effects of the LEWICE ice shape horn height on CD was similar to that 

obtained with the IRT ice shapes, i.e., ice shapes with larger horn heights produced greater drag 

increments. Moreover, the larger horns of the LS-CS22S ice shape produced a larger drag 

increment with respect to the clean wing at lower α in comparison to the LS-CS22N ice shape. 

As α was increased, the difference in drag performance between these two 22.5-min ice shapes 

diminished. 

3.2.3.3.3 Pitching-Moment Coefficient. 

From figure 3-38(e), it is observed that the pitching moments about the 25% MAC point of the 

iced wing were considerably more positive or more negative than the clean wing, depending on 

α. This was the result of a shift in the center of the aerodynamic load due to flow separation. In 

general, flow separation downstream of the iced shape resulted in a front-loaded pressure 

distribution caused by the leading-edge separation bubbles. Another contributing factor to the 

observed increase in iced wing pitching moment was the aerodynamic load on the ice shape, 

which increased as α was increased. The largest increment in CM was obtained with the LSCS22N ice shape. 

Pitching moment for the LS-SC5 configuration was comparable to the clean wing, particularly 

for α in the range of -4o to approximately 12°. As α was increased above 12o, however, the LSSC5 ice shape produced more positive or more negative CM than the clean case. The similarities 

and differences in the CM curves of the clean and LS-SC5 cases can be explained by reviewing 

the Cp distributions shown in figures 3-39 to 3-41. For α of 4o and 8o, LS-SC5 produced a lower 

suction peak than the baseline at all three spanwise stations. However, pressure recovery 

downstream of the suction peak and the Cp distributions over the lower surface were similar to 

the clean case. As a result, iced wing CM at 4o and 8o α was comparable to the clean wing. On 

the other hand, as α was increased to 12o, the iced configuration (LS-SC5) had a greater suction 

peak than the clean case at 50% semispan and over the outboard wing stations. As α was further 

3-40
increased to 16°, while both configurations (clean and LS-SC5) experienced extensive flow 

separation near the wing tip, the LS-SC5 ice shape resulted in a separation bubble on the wing 

suction surface and produced higher leading-edge suction than the clean case, as shown in 

figures 3-39(f) and 3-40(f). Consequently, the LS-SC5 case resulted in higher CM for α greater 

than 12°. 

3.2.3.3.4 Hinge-Moment Coefficient. 

Hinge-moment coefficients for all smooth LEWICE ice shapes tested are presented in 

figure 3-38(f). Hinge-moment coefficient slope (CHα) for the clean and iced wing are compared 

in table 3-6. In general, hinge moments for the iced wing were more positive or more negative 

with respect to the clean case, particularly in regions A and B of the hinge-moment curve. The 

observed increase or decrease in hinge moment was the result of flow separation over the aileron 

due to the ice shapes. In most cases, the start of region B in the iced wing CH curve was at a 

lower α, and the slope of CH was less than the clean wing. Note that the maximum and 

minimum hinge moments for all ice shape cases were bounded by the maximum and minimum 

CH of the clean wing. 

3.2.3.3.5 Pressure Distributions. 

Cp distributions are presented in figures 3-39 to 3-41 for the 15%, 50%, and 85% semispan 

stations. Near the wing tip, extensive flow separation was observed over the wing suction 

surface for most iced configurations, as shown in figure 3-41. Considering the wing pressure 

distributions with the glaze ice shapes LS-CS2, LS-CS10, LS-CS22N, and LS-CS22S, ice shapes 

with larger horns produced longer bubbles, as was the case with their IRT ice shape counterparts. 

In general, the LS-CS22N and LS-CS22S cases had comparable Cp distributions at all stations. 

However, for 12° and 16° α, the LS-CS22S had higher pressure over the wing lower surface and 

hence generated higher lift. 

3.2.3.4 Rough LEWICE Ice Shapes. 

The effects of the rough LEWICE ice shapes on wing performance are demonstrated in 

figures 3-42 to 3-45 and in tables 3-7 to 3-9. The results showed that, in most cases, the rough 

LEWICE ice shapes resulted in greater reductions in CL,stall and αstall and larger drag increments 

compared to the smooth LEWICE ice shapes. Roughness has a significant impact on the 

location of flow separation, particularly for large glaze ice shapes where the flow at high 

Reynolds numbers usually separates near the horn tip. The size of the separated region 

downstream of the ice shape is a function of the location of the separation point. Rough and 

smooth ice shapes can result in considerably different aerodynamic effects, particularly when the 

radius of the horn tips is large. For small ice shapes that do not exhibit horn features such as 

rime ice shapes, roughness may actually increase aerodynamic performance, depending on the 

Reynolds number. In this case, the location of flow separation is affected by the state of the 

boundary layer. At low Reynolds numbers, roughness can increase the energy levels in the 

boundary layer and delay flow separation, resulting in improved aerodynamic performance such 

as higher CL,stall and αstall. 

3-41
3.2.3.4.1 Lift Coefficient. 

Note that effects of higher Reynolds numbers representative of flight on the clean wing are not 

known. In most cases, the rough LEWICE ice shapes decreased CL throughout the α-range, as 

demonstrated in figures 3-42(a) and 3-42(b). One exception to this observation was the 5-min 

rime ice shape LR-SC5 that resulted in improved lift performance with respect to the clean wing. 

Tables 3-7 and 3-8 provide lift coefficients for selected angles of attack for the clean and iced 

wing cases. The reduction in CL,stall was 30%, 25%, -12.6% (increase), 17%, 53%, 61%, and 

30% for the LR-CS10, LR-IS10, LR-SC5, LR-CS2, LR-CS22N, LR-CS22S, and LR-IPSF22, 

respectively. The reduction in αstall for the glaze-iced configurations ranged from 0.1° to 3.5°,

depending on ice shape. The 5-min rime ice shape increased the clean wing αstall by 2°. With 

the exception of the 5-min rime and 2-min glaze ice shapes, all other rough LEWICE ice shapes 

reduced the linear lift slope of the clean wing. 

Observe that ice shapes with the larger horn size resulted in greater penalties in lift performance, 

i.e., LR-CS2 ice shape had higher αstall and CL,stall than LR-CS10, LR-CS22N, and LR-CS22S 

cases. Effects of horn height on CL were similar to the trend of IRT ice shapes. 

3.2.3.4.2 Drag Coefficient. 

Table 3-7 lists CD of clean and iced configurations at angles of attack of 3°, 5°, 10°, and 15°,

while table 3-8 shows the effect of rough LEWICE ice shapes on CD,min. Graphically, the CD and 

ΔCD curves are illustrated in figures 3-42(c) and 3-42(d), respectively. All roughened LEWICE 

ice shapes increased CD throughout the α-range. The increase in CD,min with respect to the clean 

case ranged from 116.7% for LR-SC5 to 3316.7% for LR-CS22S. In terms of the horn height 

effects on CD, the trends were similar to that obtained with the IRT ice shapes, i.e., ice shapes 

with larger horn heights produced greater drag increments. Moreover, the larger horns of LSCS22S ice shape produced greater drag at lower α in comparison to the LS-CS22N ice shape. 

But, as α increased, the difference in drag performance diminished. 

3.2.3.4.3 Pitching-Moment Coefficient. 

The trends in CM for the rough LEWICE ice shapes were similar to that obtained with the smooth 

LEWICE ice shapes, as demonstrated in figure 3-42(e). For positive α less than about 15°, flow 

separation downstream of the ice shapes caused the lift center to move upstream of the 25% 

MAC location, which led to more positive (leading edge up) pitching moment than the clean 

wing. However, for greater α, most of the iced wing cases experienced a large change, from 

positive to negative, in pitching moment. This was attributed to the flat pressure distribution 

associated with massive flow separation over the suction surface of the wing. For the glaze ice 

shapes LR-CS22N and LR-CS22S, the pitching moment remained positive throughout the 

positive α-range. For these two large ice shapes, extensive flow separation was present over 

both upper and lower wing surfaces for practically all positive α and the lift center did not move 

upstream of the 25% MAC location. 

3-42
3.2.3.4.4 Hinge-Moment Coefficient. 

Aileron hinge moments for the iced wing configurations are depicted in figure 3-42(f). Table 3-9 

provides the slopes and α-range for regions A, B, and C of the CH curves. The results show that 

the LR-SC5 ice shape had a CH curve similar to the clean case for practically the complete range 

of angles of attack. The ice shapes LR-CS22N and LR-CS22S resulted in large changes from the 

clean wing aileron hinge moments. In fact, for α in the range of 6° to 8° for the LR-CS22N case 

and 6° to 11° for the LR-CS22S ice shape, the aileron hinge moment was nearly zero, indicating 

zero stick force and possibly aileron float. Force reversal was the result of nearly equal pressure 

distributions over the suction and pressure sides of the aileron surfaces. The other four rough 

LEWICE ice shapes (LR-CS10, LR-IS10, LR-CS2, and LR-IPSF22) cause more positive or 

more negative hinge moments with respect to the clean wing over regions A and B of the CH

curve. Once again, the iced wing hinge moments were bounded by the clean wing CH limits. 

One exception was the 5-min rime ice shape that resulted in slightly higher hinge moments with 

respect to the clean wing in region C of the CH curve for positive α. 

3.2.3.4.5 Pressure Distributions. 

Effects of rough LEWICE ice shapes on pressure distributions are presented in figures 3-43 to 

3-45 for the 15%, 50%, and 85% semispan stations. All glaze ice shapes produced separation 

bubbles downstream of the ice horns, which reduced the suction peak near the leading edge of 

the clean wing. Among all glaze ice cases, LR-CS2 caused the least penalties on aerodynamic 

performance due to the fact that its ice horns were smaller. Consequently, for positive α, the 

separation bubble on the suction wing surface had shorter extent. Pressure distributions aft of the 

bubble and on the pressure surface were comparable to the clean wing. As the ice shape horn 

height increased (such as LR-CS10, LR-CS22N, and LR-CS22S configurations), the size of the 

separation bubble was increased. 

3.2.4 Comparison of IRT and LEWICE Ice Shapes. 

Figures 2-33 to 2-38 demonstrate that the differences in the geometric features of the LEWICE 

and IRT ice shape sections were considerable for the CS10, IS10, CS22, and IPSF22 cases. 

These differences included overall shape of ice section, height of the upper and lower ice horns, 

horn angle with respect to the horizontal and location of horn tip, and root with respect to the 

leading edge of the wing. For example, at stations A and B, the upper and lower horn tips of 

IRT-CS10 were further upstream of the LEWICE ice shape (LS-CS10) horn tips, as 

demonstrated in figure 2-33. In addition, at these two stations, the horn angles with respect to 

the horizontal were, in most cases, greater for the LEWICE ice shape. Note, however, that the 

upper and lower horn heights of the LEWICE and the IRT ice shape sections were similar for 

both spanwise stations. At station C, the LEWICE ice shape upper and lower horns extended 

further upstream and downstream with respect to the horns of the IRT casting. In addition, the 

height (distance of horn tip from airfoil LE) of the LEWICE horns was greater. Other significant 

differences between the simulated and the IRT ice shapes included the intricate roughness and 

feather-like features seen on the surface of the IRT ice shape castings presented in figure 2-33. 

To assess the effect of these geometric differences in the actual and simulated ice shapes on 

aerodynamic performance, experimental data obtained with the castings and the corresponding 

smooth and rough LEWICE ice shapes are compared in figures 3-46 to 3-69. 

3-43
 

(a) Lift coefficient (b) ΔCL vs α

(c) Drag coefficient (d) ΔCD vs α

(e) Pitching-moment coefficient (f) Hinge-moment coefficient 

FIGURE 3-46. EFFECT OF IRT-CS10, LS-CS10, AND LR-CS10 ICE SHAPES 

(ICING CONDITION 1) ON CL, CD, CM, AND CH; Re = 1.8x106; δA = 0°

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Lift coefficient

Clean (R006)

IRT-CS10 (R018)

LS-CS10 (R334)

LR-CS10 (R363)

-10 -5 0 5 10 15 20 2

Angle of attack (deg)

5

-80

-70

-60

-50

-40

-30

-20

-10

0

10

20

ΔCL (%)

IRT-CS10 (R018)

LS-CS10 (R334)

LR-CS10 (R363)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

0.00

0.05

0.10

0.15

0.20

0.25

0.30

0.35

Drag coefficient

Clean (R006)

IRT-CS10 (R018)

LS-CS10 (R334)

LR-CS10 (R363)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

0

200

400

600

800

1000

1200

1400

ΔCD (%)

IRT-CS10 (R018)

LS-CS10 (R334)

LR-CS10 (R363)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.12

-0.10

-0.08

-0.06

-0.04

-0.02

0.00

0.02

0.04

0.06

0.08

Hinge moment coefficient

Clean (R006)

IRT-CS10 (R018)

LS-CS10 (R334)

LR-CS10 (R363)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.10

-0.08

-0.06

-0.04

-0.02

0.00

0.02

0.04

0.06

0.08

0.10

Pitching moment coefficient

Clean (R006)

IRT-CS10 (R018)

LS-CS10 (R334)

LR-CS10 (R363)

3-44
(a) Lift coefficient (b) ΔCL vs α

(c) Drag coefficient (d) ΔCD vs α

(e) Pitching-moment coefficient (f) Hinge-moment coefficient 

FIGURE 3-47. EFFECT OF IRT-IS10, LS-IS10, AND LR-IS10 ICE SHAPES 

(ICING CONDITION 2) ON CL, CD, CM, AND CH; Re = 1.8x106

; δA = 0°

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Lift coefficient

Clean (R006)

IRT-IS10 (R110)

LS-IS10 (R172)

LR-IS10 (R240)

-10 -5 0 5 10 15 20 2

Angle of attack (deg)

5

-50

-40

-30

-20

-10

0

10

20

30

ΔCL (%)

IRT-IS10 (R110)

LS-IS10 (R172)

LR-IS10 (R240)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

0.00

0.05

0.10

0.15

0.20

0.25

0.30

0.35

Drag coefficient

Clean (R006)

IRT-IS10 (R110)

LS-IS10 (R172)

LR-IS10 (R240)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

0

100

200

300

400

500

600

700

800

ΔCD (%)

IRT-IS10 (R110)

LS-IS10 (R172)

LR-IS10 (R240)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.10

-0.08

-0.06

-0.04

-0.02

0.00

0.02

0.04

0.06

0.08

0.10

Pitching moment coefficient

Clean (R006)

IRT-IS10 (R110)

LS-IS10 (R172)

LR-IS10 (R240)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.12

-0.10

-0.08

-0.06

-0.04

-0.02

0.00

0.02

0.04

0.06

0.08

Hinge moment coefficient

Clean (R006)

IRT-IS10 (R110)

LS-IS10 (R172)

LR-IS10 (R240)

3-45
 

(a) Lift coefficient (b) ΔCL vs α

(c) Drag coefficient (d) ΔCD vs α

(e) Pitching-moment coefficient (f) Hinge-moment coefficient 

FIGURE 3-48. EFFECT OF IRT-SC5, LS-SC5, AND LR-SC5 ICE SHAPES 

(ICING CONDITION 3) ON CL, CD, CM, AND CH; Re = 1.8x106; δA = 0°

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Lift coefficient

Clean (R006)

IRT-SC5 (R055)

LS-SC5 (R187)

LR-SC5 (R216)

-10 -5 0 5 10 15 20 2

Angle of attack (deg)

5

-20

-15

-10

-5

0

5

10

15

ΔCL (%)

IRT-SC5 (R055)

LS-SC5 (R187)

LR-SC5 (R216)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

0.00

0.05

0.10

0.15

0.20

0.25

0.30

Drag coefficient

Clean (R006)

IRT-SC5 (R055)

LS-SC5 (R187)

LR-SC5 (R216)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-40

-20

0

20

40

60

80

100

120

140

160

ΔCD (%)

IRT-SC5 (R055)

LS-SC5 (R187)

LR-SC5 (R216)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.10

-0.08

-0.06

-0.04

-0.02

0.00

0.02

0.04

0.06

0.08

0.10

Pitching moment coefficient

Clean (R006)

IRT-SC5 (R055)

LS-SC5 (R187)

LR-SC5 (R216)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.12

-0.10

-0.08

-0.06

-0.04

-0.02

0.00

0.02

0.04

0.06

0.08

Hinge moment coefficient

Clean (R006)

IRT-SC5 (R055)

LS-SC5 (R187)

LR-SC5 (R216)

3-46
 

(a) Lift coefficient (b) ΔCL vs α

(c) Drag coefficient (d) ΔCD vs α

(e) Pitching-moment coefficient (f) Hinge-moment coefficient 

FIGURE 3-49. EFFECT OF IRT-CS2, LS-CS2, AND LR-CS2 ICE SHAPES 

(ICING CONDITION 4) ON CL, CD, CM, AND CH; Re = 1.8x106; δA = 0°

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Lift coefficient

Clean (R006)

IRT-CS2 (R070)

LS-CS2 (R206)

LR-CS2 (R232)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-25

-20

-15

-10

-5

0

5

10

15

ΔCL (%)

IRT-CS2 (R070)

LS-CS2 (R206)

LR-CS2 (R232)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

0.00

0.05

0.10

0.15

0.20

0.25

0.30

Drag coefficient

Clean (R006)

IRT-CS2 (R070)

LS-CS2 (R206)

LR-CS2 (R232)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-50

0

50

100

150

200

250

300

ΔCD (%)

IRT-CS2 (R070)

LS-CS2 (R206)

LR-CS2 (R232)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.10

-0.08

-0.06

-0.04

-0.02

0.00

0.02

0.04

0.06

0.08

0.10

Pitching moment coefficient

Clean (R006)

IRT-CS2 (R070)

LS-CS2 (R206)

LR-CS2 (R232)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.12

-0.10

-0.08

-0.06

-0.04

-0.02

0.00

0.02

0.04

0.06

0.08

Hinge moment coefficient

Clean (R006)

IRT-CS2 (R070)

LS-CS2 (R206)

LR-CS2 (R232)

3-47
 

(a) Lift coefficient (b) ΔCL vs α

(c) Drag coefficient (d) ΔCD vs α

(e) Pitching-moment coefficient

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-120

-100

-80

-60

-40

-20

0

20

40

60

ΔCL (%)

IRT-CS22 (R047)

LS-CS22N (R414)

LS-CS22S (R442)

LR-CS22N (R459)

LR-CS22S (R468)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Lift coefficient

Clean (R006)

IRT-CS22 (R047)

LS-CS22N (R414)

LS-CS22S (R442)

LR-CS22N (R459)

LR-CS22S (R468)

(f) Hinge-moment coefficient 

FIGURE 3-50. EFFECT OF IRT-CS22, LS-CS22N, LS-CS22S, LR-CS22N, AND LR-CS22S 

ICE SHAPES (ICING CONDITION 5) ON CL, CD, CM, AND CH; Re = 1.8x106

; δA = 0°

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

0.00

0.05

0.10

0.15

0.20

0.25

0.30

0.35

0.40

0.45

Drag coefficient

Clean (R006)

IRT-CS22 (R047)

LS-CS22N (R414)

LS-CS22S (R442)

LR-CS22N (R459)

LR-CS22S (R468)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

0

500

1000

1500

2000

2500

3000

3500

4000

ΔCD (%)

IRT-CS22 (R047)

LS-CS22N (R414)

LS-CS22S (R442)

LR-CS22N (R459)

LR-CS22S (R468)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.10

-0.08

-0.06

-0.04

-0.02

0.00

0.02

0.04

0.06

0.08

0.10

Pitching moment coefficient

Clean (R006)

IRT-CS22 (R047)

LS-CS22N (R414)

LS-CS22S (R442)

LR-CS22N (R459)

LR-CS22S (R468)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.12

-0.10

-0.08

-0.06

-0.04

-0.02

0.00

0.02

0.04

0.06

0.08

Hinge moment coefficient

Clean (R006)

IRT-CS22 (R047)

LS-CS22N (R414)

LS-CS22S (R442)

LR-CS22N (R459)

LR-CS22S (R468)

3-48
 

(a) Lift coefficient (b) ΔCL vs α

(c) Drag coefficient (d) ΔCD vs α

(e) Pitching-moment coefficient (f) Hinge-moment coefficient 

FIGURE 3-51. EFFECT OF IRT-IPSF22, LS-IPSF22, AND LR-IPSF22 ICE SHAPES 

(ICING CONDITION 6) ON CL, CD, CM, AND CH; Re = 1.8x106

; δA = 0°

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Lift coefficient

Clean (R006)

IRT-IPSF22 (R091)

LS-IPSF22 (R348)

LR-IPSF22 (R374)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-80

-60

-40

-20

0

20

40

60

ΔCL (%)

IRT-IPSF22 (R091)

LS-IPSF22 (R348)

LR-IPSF22 (R374)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

0.00

0.05

0.10

0.15

0.20

0.25

0.30

0.35

Drag coefficient

Clean (R006)

IRT-IPSF22 (R091)

LS-IPSF22 (R348)

LR-IPSF22 (R374)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-200

0

200

400

600

800

1000

1200

1400

ΔCD (%)

IRT-IPSF22 (R091)

LS-IPSF22 (R348)

LR-IPSF22 (R374)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.10

-0.08

-0.06

-0.04

-0.02

0.00

0.02

0.04

0.06

0.08

0.10

Pitching moment coefficient

Clean (R006)

IRT-IPSF22 (R091)

LS-IPSF22 (R348)

LR-IPSF22 (R374)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.12

-0.10

-0.08

-0.06

-0.04

-0.02

0.00

0.02

0.04

0.06

0.08

Hinge moment coefficient

Clean (R006)

IRT-IPSF22 (R091)

LS-IPSF22 (R348)

LR-IPSF22 (R374)

3-49
0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-CS10 (R018)

LS-CS10 (R334)

LR-CS10 (R363)

Clean: min Cp = -2.3254 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

IRT-CS10 (R018)

LS-CS10 (R334)

LR-CS10 (R363)

Clean: max Cp = 0.7099 at x/c = 0

(a) Cp vs x/c (α = -4°) (b) Cp vs x/c (α = 0°) 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-CS10 (R018)

LS-CS10 (R334)

LR-CS10 (R363)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

IRT-CS10 (R018)

LS-CS10 (R334)

LR-CS10 (R363)

Clean: min Cp = -3.2976 at x/c = 0.005

(c) Cp vs x/c (α = 4°) (d) Cp vs x/c (α = 8°) 

(e) Cp vs x/c (α = 12°) (f) Cp vs x/c (α = 16°) 

FIGURE 3-52. EFFECT OF IRT-CS10, LS-CS10, AND LR-CS10 ICE SHAPES (ICING 

CONDITION 1) ON PRESSURE DISTRIBUTIONS AT 15% SEMISPAN; 

Re = 1.8x106

; δA = 0°

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

-2.5

-3.0

Pressure coefficient

Clean (R006)

IRT-CS10 (R018)

LS-CS10 (R334)

LR-CS10 (R363)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

IRT-CS10 (R018)

LS-CS10 (R334)

LR-CS10 (R363)

3-50
0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-CS10 (R018)

LS-CS10 (R334)

LR-CS10 (R363)

Clean: min Cp = -3.1817 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

IRT-CS10 (R018)

LS-CS10 (R334)

LR-CS10 (R363)

Clean: max Cp = 0.677 at x/c = 0.002

(a) Cp vs x/c (α = -4°) (b) Cp vs x/c (α = 0°) 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-CS10 (R018)

LS-CS10 (R334)

LR-CS10 (R363)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

IRT-CS10 (R018)

LS-CS10 (R334)

LR-CS10 (R363)

Clean: min Cp = -3.4743 at x/c = 0.005

(c) Cp vs x/c (α = 4°) (d) Cp vs x/c (α = 8°) 

(e) Cp vs x/c (α = 12°) (f) Cp vs x/c (α = 16°) 

FIGURE 3-53. EFFECT OF IRT-CS10, LS-CS10, AND LR-CS10 ICE SHAPES (ICING 

CONDITION 1) ON PRESSURE DISTRIBUTIONS AT 50% SEMISPAN; 

Re = 1.8x106

; δA = 0°

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-CS10 (R018)

LS-CS10 (R334)

LR-CS10 (R363)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

IRT-CS10 (R018)

LS-CS10 (R334)

LR-CS10 (R363)

3-51
0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-CS10 (R018)

LS-CS10 (R334)

LR-CS10 (R363)

Clean: max Cp = 0.6244 at x/c = 0.004

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-CS10 (R018)

LS-CS10 (R334)

LR-CS10 (R363)

Clean: min Cp = -3.5329 at x/c = 0.005

(a) Cp vs x/c (α = -4°) (b) Cp vs x/c (α = 0°) 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

IRT-CS10 (R018)

LS-CS10 (R334)

LR-CS10 (R363)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

IRT-CS10 (R018)

LS-CS10 (R334)

LR-CS10 (R363)

Clean: min Cp = -2.7 at x/c = 0.004

(c) Cp vs x/c (α = 4°) (d) Cp vs x/c (α = 8°) 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-CS10 (R018)

LS-CS10 (R334)

LR-CS10 (R363)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

IRT-CS10 (R018)

LS-CS10 (R334)

LR-CS10 (R363)

(e) Cp vs x/c (α = 12°) (f) Cp vs x/c (α = 16°) 

FIGURE 3-54. EFFECT OF IRT-CS10, LS-CS10, AND LR-CS10 ICE SHAPES (ICING 

CONDITION 1) ON PRESSURE DISTRIBUTIONS AT 85% SEMISPAN; 

Re = 1.8x106

; δA = 0°

3-52
(a) Cp vs x/c (α = -4°) (b) Cp vs x/c (α = 0°) 

(c) Cp vs x/c (α = 4°) (d) Cp vs x/c (α = 8°) 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-IS10 (R110)

LS-IS10 (R172)

LR-IS10 (R240)

Clean: min Cp = -2.3254 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

IRT-IS10 (R110)

LS-IS10 (R172)

LR-IS10 (R240)

Clean: max Cp = 0.7099 at x/c = 0

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-IS10 (R110)

LS-IS10 (R172)

LR-IS10 (R240)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

IRT-IS10 (R110)

LS-IS10 (R172)

LR-IS10 (R240)

Clean: min Cp = -3.2976 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

-2.5

-3.0

Pressure coefficient

Clean (R006)

IRT-IS10 (R110)

LS-IS10 (R172)

LR-IS10 (R240)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

IRT-IS10 (R110)

LS-IS10 (R172)

LR-IS10 (R240)

(e) Cp vs x/c (α = 12°) (f) Cp vs x/c (α = 16°) 

FIGURE 3-55. EFFECT OF IRT-IS10, LS-IS10, AND LR-IS10 ICE SHAPES (ICING 

CONDITION 2) ON PRESSURE DISTRIBUTIONS AT 15% SEMISPAN; 

Re = 1.8x106

; δA = 0°

3-53
(a) Cp vs x/c (α = -4°) (b) Cp vs x/c (α = 0°) 

(c) Cp vs x/c (α = 4°) (d) Cp vs x/c (α = 8°) 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-IS10 (R110)

LS-IS10 (R172)

LR-IS10 (R240)

Clean: min Cp = -3.1817 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

IRT-IS10 (R110)

LS-IS10 (R172)

LR-IS10 (R240)

Clean: max Cp = 0.677 at x/c = 0.002

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-IS10 (R110)

LS-IS10 (R172)

LR-IS10 (R240)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

IRT-IS10 (R110)

LS-IS10 (R172)

LR-IS10 (R240)

Clean: min Cp = -3.4743 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-IS10 (R110)

LS-IS10 (R172)

LR-IS10 (R240)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

IRT-IS10 (R110)

LS-IS10 (R172)

LR-IS10 (R240)

(e) Cp vs x/c (α = 12°) (f) Cp vs x/c (α = 16°) 

FIGURE 3-56. EFFECT OF IRT-IS10, LS-IS10, AND LR-IS10 ICE SHAPES (ICING 

CONDITION 2) ON PRESSURE DISTRIBUTIONS AT 50% SEMISPAN; 

Re = 1.8x106

; δA = 0°

3-54
(a) Cp vs x/c (α = -4°) (b) Cp vs x/c (α = 0°) 

(c) Cp vs x/c (α = 4°) (d) Cp vs x/c (α = 8°) 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-IS10 (R110)

LS-IS10 (R172)

LR-IS10 (R240)

Clean: min Cp = -3.5329 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-IS10 (R110)

LS-IS10 (R172)

LR-IS10 (R240)

Clean: max Cp = 0.6244 at x/c = 0.004

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

IRT-IS10 (R110)

LS-IS10 (R172)

LR-IS10 (R240)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

IRT-IS10 (R110)

LS-IS10 (R172)

LR-IS10 (R240)

Clean: min Cp = -2.7 at x/c = 0.004

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-IS10 (R110)

LS-IS10 (R172)

LR-IS10 (R240)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

IRT-IS10 (R110)

LS-IS10 (R172)

LR-IS10 (R240)

(e) Cp vs x/c (α = 12°) (f) Cp vs x/c (α = 16°) 

FIGURE 3-57. EFFECT OF IRT-IS10, LS-IS10, AND LR-IS10 ICE SHAPES (ICING 

CONDITION 2) ON PRESSURE DISTRIBUTIONS AT 85% SEMISPAN; 

Re = 1.8x106

; δA = 0°

3-55
0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-SC5 (R055)

LS-SC5 (R187)

LR-SC5 (R216)

Clean: min Cp = -2.3254 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

IRT-SC5 (R055)

LS-SC5 (R187)

LR-SC5 (R216)

Clean: max Cp = 0.7099 at x/c = 0

(a) Cp vs x/c (α = -4°) (b) Cp vs x/c (α = 0°) 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-SC5 (R055)

LS-SC5 (R187)

LR-SC5 (R216)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

IRT-SC5 (R055)

LS-SC5 (R187)

LR-SC5 (R216)

Clean: min Cp = -3.2976 at x/c = 0.005

(b) Cp vs x/c (α = 4°) (d) Cp vs x/c (α = 8°) 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

-2.5

-3.0

Pressure coefficient

Clean (R006)

IRT-SC5 (R055)

LS-SC5 (R187)

LR-SC5 (R216)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

-2.5

Pressure coefficient

Clean (R006)

IRT-SC5 (R055)

LS-SC5 (R187)

LR-SC5 (R216)

(e) Cp vs x/c (α = 12°) (f) Cp vs x/c (α = 16°) 

FIGURE 3-58. EFFECT OF IRT-SC5, LS-SC5, AND LR-SC5 ICE SHAPES (ICING 

CONDITION 3) ON PRESSURE DISTRIBUTIONS AT 15% SEMISPAN; 

Re = 1.8x106

; δA = 0°

3-56
0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-SC5 (R055)

LS-SC5 (R187)

LR-SC5 (R216)

Clean: min Cp = -3.1817 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

IRT-SC5 (R055)

LS-SC5 (R187)

LR-SC5 (R216)

Clean: max Cp = 0.677 at x/c = 0.002

(a) Cp vs x/c (α = -4°) (b) Cp vs x/c (α = 0°) 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-SC5 (R055)

LS-SC5 (R187)

LR-SC5 (R216)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

IRT-SC5 (R055)

LS-SC5 (R187)

LR-SC5 (R216)

Clean: min Cp = -3.4743 at x/c = 0.005

(c) Cp vs x/c (α = 4°) (d) Cp vs x/c (α = 8°) 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

IRT-SC5 (R055)

LS-SC5 (R187)

LR-SC5 (R216)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

IRT-SC5 (R055)

LS-SC5 (R187)

LR-SC5 (R216)

(e) Cp vs x/c (α = 12°) (f) Cp vs x/c (α = 16°) 

FIGURE 3-59. EFFECT OF IRT-SC5, LS-SC5, AND LR-SC5 ICE SHAPES (ICING 

CONDITION 3) ON PRESSURE DISTRIBUTIONS AT 50% SEMISPAN; 

Re = 1.8x106

; δA = 0°

3-57
0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-SC5 (R055)

LS-SC5 (R187)

LR-SC5 (R216)

Clean: min Cp = -3.5329 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-SC5 (R055)

LS-SC5 (R187)

LR-SC5 (R216)

Clean: max Cp = 0.6244 at x/c = 0.004

(a) Cp vs x/c (α = -4°) (b) Cp vs x/c (α = 0°) 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

IRT-SC5 (R055)

LS-SC5 (R187)

LR-SC5 (R216)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

IRT-SC5 (R055)

LS-SC5 (R187)

LR-SC5 (R216) Clean: min Cp = -2.7 at x/c = 0.004

(c) Cp vs x/c (α = 4°) (d) Cp vs x/c (α = 8°) 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-SC5 (R055)

LS-SC5 (R187)

LR-SC5 (R216)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

IRT-SC5 (R055)

LS-SC5 (R187)

LR-SC5 (R216)

(e) Cp vs x/c (α = 12°) (f) Cp vs x/c (α = 16°) 

FIGURE 3-60. EFFECT OF IRT-SC5, LS-SC5, AND LR-SC5 ICE SHAPES (ICING 

CONDITION 3) ON PRESSURE DISTRIBUTIONS AT 85% SEMISPAN; 

Re = 1.8x106

; δA = 0°

3-58
0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-CS2 (R070)

LS-CS2 (R206)

LR-CS2 (R232)

Clean: min Cp = -2.3254 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

IRT-CS2 (R070)

LS-CS2 (R206)

LR-CS2 (R232)

Clean: max Cp = 0.7099 at x/c = 0

(a) Cp vs x/c (α = -4°) (b) Cp vs x/c (α = 0°) 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-CS2 (R070)

LS-CS2 (R206)

LR-CS2 (R232)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

IRT-CS2 (R070)

LS-CS2 (R206)

LR-CS2 (R232)

Clean: min Cp = -3.2976 at x/c = 0.005

(c) Cp vs x/c (α = 4°) (d) Cp vs x/c (α = 8°) 

(e) Cp vs x/c (α = 12°) (f) Cp vs x/c (α = 16°) 

FIGURE 3-61. EFFECT OF IRT-CS2, LS-CS2, AND LR-CS2 ICE SHAPES (ICING 

CONDITION 4) ON PRESSURE DISTRIBUTIONS AT 15% SEMISPAN; 

Re = 1.8x106

; δA = 0°

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

IRT-CS2 (R070)

LS-CS2 (R206)

LR-CS2 (R232)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

-2.5

-3.0

Pressure coefficient

Clean (R006)

IRT-CS2 (R070)

LS-CS2 (R206)

LR-CS2 (R232)

3-59
0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-CS2 (R070)

LS-CS2 (R206)

LR-CS2 (R232)

Clean: min Cp = -3.1817 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

IRT-CS2 (R070)

LS-CS2 (R206)

LR-CS2 (R232)

Clean: max Cp = 0.677 at x/c = 0.002

(a) Cp vs x/c (α = -4°) (b) Cp vs x/c (α = 0°) 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-CS2 (R070)

LS-CS2 (R206)

LR-CS2 (R232)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

IRT-CS2 (R070)

LS-CS2 (R206)

LR-CS2 (R232)

Clean: min Cp = -3.4743 at x/c = 0.005

(c) Cp vs x/c (α = 4°) (d) Cp vs x/c (α = 8°) 

(e) Cp vs x/c (α = 12°) (f) Cp vs x/c (α = 16°) 

FIGURE 3-62. EFFECT OF IRT-CS2, LS-CS2, AND LR-CS2 ICE SHAPES (ICING 

CONDITION 4) ON PRESSURE DISTRIBUTIONS AT 50% SEMISPAN; 

Re = 1.8x106

; δA = 0°

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

IRT-CS2 (R070)

LS-CS2 (R206)

LR-CS2 (R232)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-CS2 (R070)

LS-CS2 (R206)

LR-CS2 (R232)

3-60
0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-CS2 (R070)

LS-CS2 (R206)

LR-CS2 (R232)

Clean: min Cp = -3.5329 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-CS2 (R070)

LS-CS2 (R206)

LR-CS2 (R232)

Clean: max Cp = 0.6244 at x/c = 0.004

(a) Cp vs x/c (α = -4°) (b) Cp vs x/c (α = 0°) 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

IRT-CS2 (R070)

LS-CS2 (R206)

LR-CS2 (R232)

Clean: min Cp = -2.7 at x/c = 0.004

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-CS2 (R070)

LS-CS2 (R206)

LR-CS2 (R232)

(c) Cp vs x/c (α = 4°) (d) Cp vs x/c (α = 8°) 

(e) Cp vs x/c (α = 12°) (f) Cp vs x/c (α = 16°) 

FIGURE 3-63. EFFECT OF IRT-CS2, LS-CS2, AND LR-CS2 ICE SHAPES (ICING 

CONDITION 4) ON PRESSURE DISTRIBUTIONS AT 85% SEMISPAN; 

Re = 1.8x106

; δA = 0°

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-CS2 (R070)

LS-CS2 (R206)

LR-CS2 (R232)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

IRT-CS2 (R070)

LS-CS2 (R206)

LR-CS2 (R232)

3-61
(a) Cp vs x/c (α = -4°) (b) Cp vs x/c (α = 0°) 

(c) Cp vs x/c (α = 4°) 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-CS22 (R047)

LS-CS22N (R414)

LS-CS22S (R442)

LR-CS22N (R459)

LR-CS22S (R468)

Clean: min Cp = -2.3254 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-CS22 (R047)

LS-CS22N (R414)

LS-CS22S (R442)

LR-CS22N (R459)

LR-CS22S (R468)

Clean: max Cp = 0.7099 at x/c = 0

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-CS22 (R047)

LS-CS22N (R414)

LS-CS22S (R442)

LR-CS22N (R459)

LR-CS22S (R468)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

IRT-CS22 (R047)

LS-CS22N (R414)

LS-CS22S (R442)

LR-CS22N (R459)

LR-CS22S (R468)

Clean: min Cp = -3.2976 at x/c = 0.005

(d) Cp vs x/c (α = 8°) 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

-2.5

-3.0

Pressure coefficient

Clean (R006)

IRT-CS22 (R047)

LS-CS22N (R414)

LS-CS22S (R442)

LR-CS22N (R459)

LR-CS22S (R468)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

IRT-CS22 (R047)

LS-CS22N (R414)

LS-CS22S (R442)

LR-CS22N (R459)

LR-CS22S (R468)

(e) Cp vs x/c (α = 12°) (f) Cp vs x/c (α = 16°) 

FIGURE 3-64. EFFECT OF IRT-CS22, LS-CS22N, LS-CS22S, LR-CS22N, AND LR-CS22S 

ICE SHAPES (ICING CONDITION 5) ON PRESSURE DISTRIBUTIONS AT 15% 

SEMISPAN; Re = 1.8x106

; δA = 0°

3-62
(a) Cp vs x/c (α = -4°) (b) Cp vs x/c (α = 0°) 

(c) Cp vs x/c (α = 4°) 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-CS22 (R047)

LS-CS22N (R414)

LS-CS22S (R442)

LR-CS22N (R459)

LR-CS22S (R468)

Clean: min Cp = -3.1817 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

IRT-CS22 (R047)

LS-CS22N (R414)

LS-CS22S (R442)

LR-CS22N (R459)

LR-CS22S (R468)

Clean: max Cp = 0.677 at x/c = 0.002

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-CS22 (R047)

LS-CS22N (R414)

LS-CS22S (R442)

LR-CS22N (R459)

LR-CS22S (R468)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

IRT-CS22 (R047)

LS-CS22N (R414)

LS-CS22S (R442)

LR-CS22N (R459)

LR-CS22S (R468)

Clean: min Cp = -3.4743 at x/c = 0.005

(d) Cp vs x/c (α = 8°) 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-CS22 (R047)

LS-CS22N (R414)

LS-CS22S (R442)

LR-CS22N (R459)

LR-CS22S (R468)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

IRT-CS22 (R047)

LS-CS22N (R414)

LS-CS22S (R442)

LR-CS22N (R459)

LR-CS22S (R468)

(e) Cp vs x/c (α = 12°) (f) Cp vs x/c (α = 16°) 

FIGURE 3-65. EFFECT OF IRT-CS22, LS-CS22N, LS-CS22S, LR-CS22N, AND LR-CS22S 

ICE SHAPES (ICING CONDITION 5) ON PRESSURE DISTRIBUTIONS AT 50% 

SEMISPAN; Re = 1.8x106; δA = 0°

3-63
(a) Cp vs x/c (α = -4°) (b) Cp vs x/c (α = 0°) 

(c) Cp vs x/c (α = 4°) 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-CS22 (R047)

LS-CS22N (R414)

LS-CS22S (R442)

LR-CS22N (R459)

LR-CS22S (R468)

Clean: max Cp = 0.6244 at x/c = 0.004

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-CS22 (R047)

LS-CS22N (R414)

LS-CS22S (R442)

LR-CS22N (R459)

LR-CS22S (R468)

Clean: min Cp = -3.5329 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

IRT-CS22 (R047)

LS-CS22N (R414)

LS-CS22S (R442)

LR-CS22N (R459)

LR-CS22S (R468)

Clean: min Cp = -2.7 at x/c = 0.004

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

IRT-CS22 (R047)

LS-CS22N (R414)

LS-CS22S (R442)

LR-CS22N (R459)

LR-CS22S (R468)

(d) Cp vs x/c (α = 8°) 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-CS22 (R047)

LS-CS22N (R414)

LS-CS22S (R442)

LR-CS22N (R459)

LR-CS22S (R468)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

IRT-CS22 (R047)

LS-CS22N (R414)

LS-CS22S (R442)

LR-CS22N (R459)

LR-CS22S (R468)

(e) Cp vs x/c (α = 12°) (f) Cp vs x/c (α = 16°) 

FIGURE 3-66. EFFECT OF IRT-CS22, LS-CS22N, LS-CS22S, LR-CS22N, AND LR-CS22S 

ICE SHAPES (ICING CONDITION 5) ON PRESSURE DISTRIBUTIONS AT 85% 

SEMISPAN; Re = 1.8x106; δA = 0°

3-64
0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-IPSF22 (R091)

LS-IPSF22 (R348)

LR-IPSF22 (R374)

Clean: min Cp = -2.3254 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

IRT-IPSF22 (R091)

LS-IPSF22 (R348)

LR-IPSF22 (R374)

Clean: max Cp = 0.7099 at x/c = 0

(a) Cp vs x/c (α = -4°) (b) Cp vs x/c (α = 0°) 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-IPSF22 (R091)

LS-IPSF22 (R348)

LR-IPSF22 (R374)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

IRT-IPSF22 (R091)

LS-IPSF22 (R348)

LR-IPSF22 (R374)

Clean: min Cp = -3.2976 at x/c = 0.005

(c) Cp vs x/c (α = 4°) (d) Cp vs x/c (α = 8°) 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

-2.5

-3.0

Pressure coefficient

Clean (R006)

IRT-IPSF22 (R091)

LS-IPSF22 (R348)

LR-IPSF22 (R374)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

IRT-IPSF22 (R091)

LS-IPSF22 (R348)

LR-IPSF22 (R374)

(e) Cp vs x/c (α = 12°) (f) Cp vs x/c (α = 16°) 

FIGURE 3-67. EFFECT OF IRT-IPSF22, LS-IPSF22, AND LR-IPSF22 ICE SHAPES (ICING 

CONDITION 6) ON PRESSURE DISTRIBUTIONS AT 15% SEMISPAN; 

Re = 1.8x106

; δA = 0°

3-65
0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-IPSF22 (R091)

LS-IPSF22 (R348)

LR-IPSF22 (R374)

Clean: min Cp = -3.1817 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-IPSF22 (R091)

LS-IPSF22 (R348)

LR-IPSF22 (R374)

Clean: max Cp = 0.677 at x/c = 0.002

(a) Cp vs x/c (α = -4°) (b) Cp vs x/c (α = 0°) 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-IPSF22 (R091)

LS-IPSF22 (R348)

LR-IPSF22 (R374)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

IRT-IPSF22 (R091)

LS-IPSF22 (R348)

LR-IPSF22 (R374)

Clean: min Cp = -3.4743 at x/c = 0.005

(c) Cp vs x/c (α = 4°) (d) Cp vs x/c (α = 8°) 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-IPSF22 (R091)

LS-IPSF22 (R348)

LR-IPSF22 (R374)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

IRT-IPSF22 (R091)

LS-IPSF22 (R348)

LR-IPSF22 (R374)

(e) Cp vs x/c (α = 12°) (f) Cp vs x/c (α = 16°) 

FIGURE 3-68. EFFECT OF IRT-IPSF22, LS-IPSF22, AND LR-IPSF22 ICE SHAPES (ICING 

CONDITION 6) ON PRESSURE DISTRIBUTIONS AT 50% SEMISPAN; 

Re = 1.8x106

; δA = 0°

3-66
0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-IPSF22 (R091)

LS-IPSF22 (R348)

LR-IPSF22 (R374)

Clean: min Cp = -3.5329 at x/c = 0.005

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-IPSF22 (R091)

LS-IPSF22 (R348)

LR-IPSF22 (R374)

Clean: max Cp = 0.6244 at x/c = 0.004

(a) Cp vs x/c (α = -4°) (b) Cp vs x/c (α = 0°) 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

IRT-IPSF22 (R091)

LS-IPSF22 (R348)

LR-IPSF22 (R374)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

Pressure coefficient

Clean (R006)

IRT-IPSF22 (R091)

LS-IPSF22 (R348)

LR-IPSF22 (R374)

Clean: min Cp = -2.7 at x/c = 0.004

(c) Cp vs x/c (α = 4°) (d) Cp vs x/c (α = 8°) 

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

Pressure coefficient

Clean (R006)

IRT-IPSF22 (R091)

LS-IPSF22 (R348)

LR-IPSF22 (R374)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

Clean (R006)

IRT-IPSF22 (R091)

LS-IPSF22 (R348)

LR-IPSF22 (R374)

(e) Cp vs x/c (α = 12°) (f) Cp vs x/c (α = 16°) 

FIGURE 3-69. EFFECT OF IRT-IPSF22, LS-IPSF22, AND LR-IPSF22 ICE SHAPES (ICING 

CONDITION 6) ON PRESSURE DISTRIBUTIONS AT 85% SEMISPAN; 

Re = 1.8x106

; δA = 0°

3-67
3.2.4.1 Icing Condition 1. 

Lift, drag, and hinge-moment coefficients for this case are presented in figure 3-46. The results 

indicate the following: 

• Lift coefficients for the IRT, smooth, and rough LEWICE ice shapes were in good overall 

agreement for α in the range of -7° to 7°, as shown in figure 3-46(a). All ice shapes 

resulted in a lower lift slope with respect to the clean wing. Over the stall region, the lift 

curves for the rough LEWICE ice shape (LR-CS10) and that of the IRT casting (IRTCS10) were in good correlation. The smooth LEWICE ice shape (LS-CS10), however, 

resulted in higher CL,stall and αstall. The reduction in CL,stall with respect to the clean wing 

for the IRT-CS10, LS-CS10, and LR-CS10 ice shapes was 38%, 18%, and 30% 

respectively. The change in αstall with respect to the clean wing case was -24%, 5.8%, 

and -1.4% for ice shapes IRT-CS10, LS-CS10, and LR-CS10 respectively. 

• All ice shapes caused a significant increase in CD (e.g., 1100% to 1200% in CD,min) with 

respect to the clean wing, as demonstrated in tables 3-1, 3-2, 3-4, 3-5, 3-7, and 3-8. Drag 

coefficients obtained with all three ice shapes were in good overall agreement. However, 

the drag due to LEWICE ice shapes was, in general, higher than the IRT ice shape 

casting, as shown in figure 3-46(c). For α greater than 10°, the smooth LEWICE ice 

shape produced higher drag compared to the rough LEWICE shape. This was the result 

of more extensive flow separation downstream of the smooth LEWICE ice horns. In 

general, the size and extent of flow separation downstream of large glaze ice shapes are 

affected by the roughness of the horn tip. 

• Pitching-moment characteristics of clean and iced configurations are presented in 

figure 3-46(e). The results show that all three ice shapes produced pitching moments that 

were considerably higher or lower than the clean wing, depending on α. The trends in 

pitching moment with α for the ice shapes IRT-CS10, LS-CS10, and LR-CS10 were 

similar. As α was increased from -8° to αstall, pitching moment was increased from 

negative (nose down) to positive (nose up) for all iced wing cases. This was mainly due 

to the leading-edge bubbles that resulted in a forward shift (ahead of the 25% MAC 

point) of the lift vector. When the bubble was on the lower surface (negative α), lift was 

negative, and the lift vector was ahead of the 25% MAC, resulting in negative CM. 

Positive α had the opposite effect on lift and CM. Past stall, CM exhibited a downward 

trend as the load center moved aft of the 25% MAC point. This was due to massive flow 

separation over the wing, which resulted in a flat pressure distribution over the upper 

surface of the wing. Notable differences in the magnitude of the pitching moment of the 

IRT-CS10 ice shape compared to the LEWICE ice shapes were observed for α in the 

range of 6 to about 15°. These differences were, again, a function of the extent of the 

flow separation over each ice shape. 

• Aileron hinge-moment coefficients for the IRT and LEWICE ice shapes are presented in 

figure 3-46(f). In all cases, the hinge moments due to the ice shapes were within the 

range defined by the minimum and maximum hinge moments of the clean wing. In 

general, the increase in hinge moments caused by the LEWICE ice shapes was greater 

3-68
than the IRT ice shape for α in the range of -3° to 20°. Note that the CHα in region A of 

the LEWICE ice shape curves was considerably higher than the IRT ice shape. 

3.2.4.2 Icing Condition 2. 

Lift, drag, and hinge-moment coefficients for this ice shape configuration are presented in 

figure 3-47. The results indicate the following: 

• Lift coefficients for the IRT, smooth, and rough LEWICE ice shapes were in good overall 

agreement for α in the range of -7° to 12°, as shown in figure 3-47(a). The lift slope for 

all three ice shapes was the same as the clean wing. The reduction in CL,stall with respect 

to the clean wing for the IRT-IS10, LS-IS10, and LR-IS10 cases was 26.4%, 21.8% and 

25.3% respectively. The reduction in αstall with respect to the clean wing was 23%, 1.4%, 

and 1.4% for ice shapes IRT-IS10, LS-IS10, and LR-IS10 respectively. 

• Drag coefficients for all three ice shapes were in good overall agreement for α in the 

range of 6° to 20°, as shown in figure 3-47(c). For α less than 6o, the IRT-IS10 ice shape 

resulted in higher drag than that obtained with the smooth and rough LEWICE ice 

shapes. The increase in CD,min for the iced wing with respect to the clean wing case was 

683%, 500%, and 550% for the IRT-IS10, LS-IS10, and LR-IS10 ice shapes respectively. 

• Pitching-moment coefficients for the IRT-IS10, LS-IS10, and LR-IS10 ice shapes are 

shown in figure 3-47(e). For α in the linear lift range, the results show that all ice shapes 

exhibited similar trends in CM behavior. Specifically, a maximum CM, of approximately 

0.03 was attained at α of about 7°. However, beyond 7° AOA, the wing with the IRTIS10 ice shape exhibited a drop followed by a small increase in CM, indicating that the lift 

vector was moving back and forth. The LEWICE ice shapes resulted in a flat pitchingmoment curve with positive CM values for α in the range 7° to about 14°. After wing 

stall, all ice shapes manifested decreasing CM (more negative) with α. 

• Aileron hinge-moment coefficients for the IRT and LEWICE ice shapes are presented in 

figure 3-47(f). The ice shapes increased aileron hinge moments for α in the range of -6°

to 13°. In all cases, however, the hinge moments due to the ice shapes were within the 

range defined by the minimum and maximum hinge moments of the clean wing. The 

increase in the iced wing hinge moment with respect to the clean wing for the LEWICE 

cases was greater than that obtained with the IRT ice shape for α in the range of 5° to 

10°. 

3.2.4.3 Icing Condition 3. 

The IRT and LEWICE ice shapes were very similar in shape and size, as demonstrated in 

figure 2-35. Aerodynamic performance of the GLC-305 wing with the IRT and LEWICE rime 

ice shapes is presented in figure 3-48. The data presented indicate the following: 

• Note that effects of higher Reynolds numbers representative of flight on the clean wing 

are not known. The lift coefficients for the clean and all iced wing cases were in good 

3-69
overall agreement for α in the range of -8° to 12°, as shown in figure 3-48(a). The ice 

shapes tested resulted in 3.4% to 12.6% higher CL,stall and 7.2% to 14.5% higher αstall

with respect to the clean wing. The main reason for the observed increase in CL,stall and 

αstall for the iced wing cases was the wing leading edge droop (see figure 2-35) feature of 

these three ice shapes. Note that the rough LEWICE shape resulted in the largest lift 

performance increment. This was probably due to delayed boundary layer separation 

caused by the grit roughness. 

• The IRT and rough LEWICE ice shapes resulted in higher drag with respect to the clean 

wing for most of the α-range tested. The smooth LEWICE ice shape (LS-SC5) increased 

drag for α up to 10°, as shown in figure 3-48(c). At higher angles of attack, however, the 

drag of LS-SC5 was in most cases lower than that of the clean wing. The increase in the 

iced wing CD,min with respect to the clean wing case was 133%, 67%, and 117% for the 

IRT-SC5, LS-SC5, and LR-SC5 ice shapes respectively. 

• From figure 3-48(e), it is observed that CM was gradually increased with AOA prior to 

stall. But, as the IRT-SC5 and LR-SC5 configurations stalled, CM decreased 

considerably and eventually became considerably more negative compared to the clean 

wing. As for LS-SC5 ice shape, its CM distribution maintained more positive CM than 

that of the clean case for angles of attack greater than 7°. 

• Aileron hinge-moment coefficients for the IRT and LEWICE ice shapes were very 

similar to that of the clean wing for angles of attack greater than -4°, as shown in 

figure 3-48(f). Between α of -8° and -4°, the ice shapes resulted in lower aileron hingemoment coefficient. 

3.2.4.4 Icing Condition 4. 

The IRT and LEWICE ice shape sections presented in figure 2-36 were in good overall 

agreement. Note that the cross section area of LEWICE ice shape at station A was larger than 

that of the IRT ice shape casting. Aerodynamic performance degradation due to the IRT and the 

LEWICE ice shapes is presented in figure 3-49. The experimental lift, drag and, hinge-moment 

coefficients presented indicate the following: 

• The lift coefficients of the clean and iced wing cases were in good agreement over the 

linear lift range corresponding to angles of attack between -6° and 9°, as shown in 

figure 3-49(a). The reduction in the CL,stall with respect to the clean wing caused by the 

IRT, smooth LEWICE, and rough LEWICE ice shapes were 11.5%, 13.8%, and 17.2% 

respectively. The corresponding change in αstall with respect to the clean wing was -8%, 

6.5%, and -1%. Note that the rough LEWICE shape resulted in the largest lift reduction 

near stall, while the IRT casting caused the largest reduction in αstall. 

• The IRT and rough LEWICE ice shapes resulted in higher drag with respect to the clean 

wing throughout the α-range, as shown in figure 3-49(c). The increase in the iced wing 

CD,min with respect to that of the clean wing was 200%, 183%, and 233% for the IRTCS2, LS-CS2, and LR-CS2 ice shapes respectively. 

3-70
• Pitching moment trends for all three ice shapes were very similar. In general, the iced 

wing CM was more negative or more positive than that of the clean wing throughout the 

α-range. The difference between the clean and the iced wing CM ranged from -0.04 to 

0.02. 

• The aileron hinge moment behavior for the clean and IRT-CS2 cases were very similar 

for angles of attack in the range of -7° to 14°. The LEWICE ice shapes increased aileron 

hinge moments with respect to the clean case over region B of the CH curve. For all ice 

shapes, the hinge-moment coefficients were within the range defined by the hinge 

moments of the clean wing. 

3.2.4.5 Icing Condition 5. 

The IRT and LEWICE ice shape sections presented in figure 2-37 exhibited large differences in 

horn shape, size, and overall section area. In addition, the IRT-CS22 ice shape had complex 

roughness and scallop features not present in the LEWICE ice shapes. Two sets of smooth and 

rough LEWICE ice shapes were tested for this icing condition as discussed in section 2.2.3.2. 

One was based on computations using sections normal (N) to the wing leading edge while the 

other was computed using streamwise (S) sections. The wing aerodynamic performance data for 

the IRT, the smooth, and rough LEWICE ice shapes presented in figure 3-50 indicate the 

following: 

• The iced wing performance exhibited large reductions in lift, αstall and in lift slope with 

respect to that of the clean wing. The largest degradation in lift was caused by the IRTCS22 ice shape. The rough LEWICE ice shapes caused larger reductions in lift than their 

smooth counterparts. The rough LEWICE ice shape LR-CS22S obtained using 

streamwise wing sections and free-stream velocity resulted in larger lift degradation than 

that obtained with the rough LEWICE ice shape LR-CS22N, which was computed with 

the normal component of the free-stream velocity and airfoil sections normal to the wing 

leading edge. The difference in lift reduction due to the IRT-CS22 ice shape and that 

caused by the rough LEWICE ice shapes was considerable, particularly for angles of 

attack in the range of 0° to approximately 13°. The reduction in the CL,stall with respect to 

the clean wing caused by the IRT-CS22, LS-CS22N, LS-CS22S, LR-CS22N, and LRCS22S ice shapes was 93.6%, 44%, 38%, 53%, and 61% respectively. The 

corresponding reduction in αstall with respect to the clean wing was 56%, 2%, 3%, 17%, 

and 25%. 

• All ice shapes tested caused very large drag increments, as shown in figure 3-50(c). The 

increase in the CD,min for the iced wing was 3533%, 2033%, 2367%, 2933%, and 3317% 

for the ice shapes IRT-CS22, LS-CS22N, LS-CS22S, LR-CS22N, and LR-CS22S 

respectively. The LEWICE ice shape that provided the best drag correlation with the 

IRT-CS22 ice shape was the LR-CS22S. 

• Figure 3-50(e) shows pitching moment characteristics of the wing with all five CS22 ice 

shapes. Separated flow downstream of both upper and lower ice horns on the IRT-CS22 

ice shape led to small changes in CM compared to the clean case, despite the large 

differences in the wing pressure distributions obtained with the IRT-CS22 ice shape and 

3-71
the clean wing. The four LEWICE ice shapes produced significantly more negative or 

more positive CM than that of the clean wing. Also, the difference in CM,clean-CM,LEWICE

ranged from 0.07 at negative AOA to -0.085 at positive AOA. With increasing positive 

α, the separation bubbles behind the glaze ice horns shifted the load center upstream of 

the quarter-chord location and hence increased CM. Conversely, beyond αstall, separated 

flow over the upper surface and attached flow over the lower surface moved the center of 

pressure downstream and led to decreased CM. It is worth noting that both LS-CS22N 

and LR-CS22S ice shapes had higher maximum CM than their CS22S counterparts. Also, 

the break from positive to negative values in the CM curve occurred at a higher α for the 

CS22N ice shapes. 

• The aileron hinge moment behavior for the clean and iced wing cases exhibited large 

differences, as shown in figure 3-50(f). However, for all ice shapes, the hinge-moment 

coefficients were practically within the range defined by the minimum and maximum 

hinge moments of the clean wing. The LEWICE ice shape that provided the best CH

correlation with the IRT casting was the LS-CS22N. 

3.2.4.6 Icing Condition 6. 

The IRT and LEWICE ice shape sections presented in figure 2-38 demonstrated notable 

differences in size and shape. However, the horn angles and horn heights of the IRT ice shape 

castings and the LEWICE ice shapes were similar. Aerodynamic performance degradation due 

to the IRT and the LEWICE ice shapes is presented in figure 3-51. The experimental lift, drag 

and hinge-moment coefficients presented in these figures indicate the following: 

• The lift coefficient for the iced wing exhibited lower lift and linear lift slope than that of 

the clean wing, as evident in figure 3-51(a). The reduction in the CL,stall of the iced wing 

with respect to the clean wing was 39%, 21% and 30% for the IRT, smooth LEWICE, 

and rough LEWICE ice shapes respectively. The corresponding change in αstall with 

respect to the clean wing was -24%, 6%, and -9%. Note that the IRT-IPSF22 ice shape 

resulted in the largest reduction of CL,stall and αstall with respect to the clean wing. The lift 

curves for the IRT and rough LEWICE ice shapes were in good agreement except near 

stall where the IRT shape caused approximately 1% to 8% more reduction in lift with 

respect to the clean wing. 

• All ice shapes resulted in significant drag increments with respect to the clean wing drag, 

as shown in figure 3-51(c). The increase in CD,min of the iced wing with respect to the 

clean configuration was 1200%, 1000%, and 983% for the IRT-IPSF22, LS-IPSF22 and 

LR-IPSF22 ice shapes respectively. The trends in the IRT and LEWICE ice shapes drag 

curves were similar. However, the IRT ice shape caused higher drag at angles of attack 

in the range of -8° to 11°. For α greater than 11°, the LEWICE ice shapes resulted in 

higher drag. 

• The trends in CM for the IPSF22 ice shapes were quite similar to that obtained with the 

CS10 ice shapes. In general, the LS-IPSF22 and LR-IPSF22 cases produced similar 

pitching moment characteristics, as demonstrated in figure 3-51(e). The observed 

3-72
differences in CM between IRT and LEWICE ice shapes included more positive CM for 

the LEWICE ice shapes for angles of attack between 6° and 18° and higher dCM/dα after 

α of approximately 12°. The iced wing cases resulted in considerably more positive CM

than the clean wing for angles of attack in the range of 1° to 11° for the IRT-IPSF22 and 

1° to 15° for the LEWICE ice shapes. 

• The iced wing aileron hinge moment was within the range defined by the minimum and 

maximum hinge moments of the clean wing for all ice shape cases. For most angles of 

attack, the LEWICE ice shapes caused a larger increase (more negative or more positive) 

in hinge moments than that caused by the IRT ice shape. 

3.2.4.7 Summary. 

With the exception of icing condition 5, the rough LEWICE ice shapes caused similar loss in lift 

compared with the IRT ice shape castings. The difference in iced wing CL,stall (IRT CL,stall - 

rough LEWICE CL,stall) obtained with the rough LEWICE and IRT ice shapes was -0.07, -0.01, 

-0.08, 0.05, -0.35, -0.28, and -0.08 for CS10, IS10, SC5, CS2, CS22N, CS22S, and IPSF22 

respectively. In all but one case, the IRT ice shape castings resulted in higher reduction in CL,stall

compared to that obtained with the rough LEWICE ice shapes. Again, with the exception of 

icing condition 5, the observed differences in CL,stall reduction between the IRT and LEWICE ice 

shapes were in the range of 1% to 9% with respect to the clean wing CL,stall which was equal to 

0.87. In addition, the behavior of the lift curves obtained with the IRT and LEWICE ice shapes 

was very similar. 

In general, the increment in wing drag caused by the LEWICE and IRT ice shapes was in good 

correlation. The difference in iced wing CD,min (rough LEWICE CD,min - IRT CD,min) for the 

rough LEWICE and IRT cases was 0.004, -0.008, -0.001, 0.002, -0.036, -0.013, and -0.013 for 

CS10, IS10, SC5, CS2, CS22N, CS22S, and IPSF22 respectively. 

With the exception of icing condition 5, the trends in aileron hinge moments obtained with the 

rough LEWICE and IRT ice shapes were in good agreement. For icing conditions 2 and 3, the 

IRT and LEWICE CH magnitudes were in good correlation throughout the α-range. For icing 

conditions 1, 4, and 6, the LEWICE ice shapes resulted in larger hinge moments over regions A 

and B of the CH curve compared with the IRT ice shapes. For icing condition 5, the behavior of 

CH for the rough LEWICE and IRT ice shapes was considerably different. This was mainly due 

to higher suction over the wing upper surface caused by the rough LEWICE ice shapes. 

In view of the differences in the geometric features of the large glaze IRT and LEWICE ice 

shapes tested, the LEWICE ice shapes provided a good engineering approximation to the 

aerodynamic effects of the IRT ice shapes for five out of the six cases tested. 

3.2.4.8 Pressure Distributions—Clean and All Iced Configurations. 

Pressure distributions for the clean wing and all 20 ice shape configurations tested are presented 

in figures 3-52 to 3-69 for α in the range of -4° to 16° and for spanwise locations corresponding 

to 15% semispan (near wing root), 50% semispan, and 85% semispan (near wing tip). The 

3-73
pressure data provided insight into the change in wing load distribution with α and the stalling 

behavior of the clean and iced wing. In the discussion below, reference is made to three angles 

of attack, namely the wing geometric angle of attack, the local geometric angle of attack, and the 

effective angle of attack. These angles are defined as follows: 

• Geometric angle of attack (α): this is the wing angle of attack as set on the wind tunnel 

turntable and is the α used in all force and moment plots. 

• Local geometric angle of attack (αlocal): this is the geometric angle of attack at a given 

spanwise station, which can be determined from α and the local geometric twist. 

• Effective angle of attack (αe): this angle of attack is the sum of αlocal and the local 

upwash or downwash angle. Upwash or downwash is caused by the wing tip vortex and 

depends on wing lift. 

For negative α, the wing upper surface was at a higher pressure than the lower surface. Due to 

wing twist, which was 0° at the root and -4° at the tip, the αlocal varied from root to tip. Thus, at 

negative α, the wing root was at a lower αlocal than the wing tip. In other words, the αlocal

increased from root to tip. At positive α, the wing tip was at a lower αlocal than the wing root. 

However, for positive α, the wing experienced upwash, which increased as the α was increased. 

Thus, for positive lift, the wing upwash was higher near the wing tip than near the wing root. 

The term suction used in the discussion of the clean and iced wing pressures refers to a region of 

flow where the static pressure is lower than in the free stream. Suction is associated with 

negative pressure coefficients and is usually observed in regions of the wing where the 

combination of wing curvature and α accelerates the flow to speeds higher than the free stream. 

In most cases, the pressure data presented in figures 3-52 to 3-69 indicated that the addition of 

ice shapes resulted in a dramatic change in the pressure distribution mainly due to a combination 

of leading-edge separation bubbles and extensive flow separation over the wing. Flow 

separation is a function of ice shape and α. The length of the separated region increased as α, or 

the height of the ice shape horn, were increased. The large changes in surface pressures 

downstream of the ice shapes were responsible for the observed behavior of the iced wing force 

and moment coefficients. 

Separated flows contain vortex structures, bubbles, and free shear layers that have a significant 

impact on the flow field. Vortex structures may be fixed in location as in the case of cavity 

flows and large stationary bubbles, or they can convect with the flow and interact with other 

vortices, free shear layers, or wall-bounded flows. Vortices can also stretch, depending on 

external forces (normal and shear forces) imposed on the vortex. The shape and strength of 

vortices in close proximity to the airfoil surface can produce considerable changes in surface 

pressures and, therefore, in aerodynamic loads. 

The flow over iced wings, particularly for cases involving glaze ice shapes with large horns, 

features large regions of flow separation downstream of the ice shapes, which are typically 

referred to as leading-edge bubbles or long bubbles. These bubbles are regions of viscous flow 

3-74
bound by the inviscid flow streamlines. The streamlines detach from the wing surface upstream 

of the bubble and reattach downstream of the bubble. If reattachment takes place over the wing 

surface, then the bubble is closed. The location of the reattachment point is a function of α and 

horn size. As α is increased, the bubble becomes progressively longer and eventually massive 

flow separation takes place once the bubble reaches the wing TE. In some cases, a long bubble 

can combine with separated flow initiated at the TE, and massive flow separation can occur 

before the bubble reaches the wing TE. 

For swept wings, LE separation bubbles may include a combination of LE flow separation and 

LE vortex flow. When sweep is incorporated in a wing, a conical vortex lying on the wing 

surface can occur, as discussed in reference 18. This LE vortex results from both the LE 

separation bubble and the spanwise pressure gradient due to sweep. The vortex cross section 

grows nearly perpendicular to the wing LE (conical vortex) in the spanwise direction. The 

growth of vortex diameter as the wing tip is approached is the result of a combination of vortex 

flow with the thick boundary layer transported by the spanwise flow from the inboard sections. 

With increased α, the vortex becomes stronger near the inboard sections but diffuses near the 

outboard sections. Near the wing tip, the diffused conical LE vortex can combine with the wing 

tip vortex, and complex flow separation patterns are often observed. LE vortices were observed 

with both the clean and iced wing configurations tested as α was increased. 

In general, pressure distributions for wing flows with LE bubbles exhibit a considerable 

reduction in LE suction and a rounding of the LE pressure distribution. As the α is increased, 

the bubble becomes longer in the streamwise direction and the suction over the bubble is 

typically reduced. This led to a flat pressure distribution over the wing surface. In general, LE 

bubbles shift the load towards the LE. However, as the bubble becomes longer and massive flow 

separation takes place, the wing load shifts in the downstream direction. Massive flow 

separation is an unsteady phenomenon that involves convection of vortices, which cause 

significant changes to the load distribution over the wing and, thus, to wing pitching moments. 

Flow separation near the outboard portion of the iced wings could also cause large changes in the 

load over aileron controls, and in some cases, it could reduce or even eliminate the control 

effectiveness. Flow separation over the aileron is also responsible for control force reversal and 

potential large changes in control stick forces. Thus, bubble formation, growth, and bursting can 

cause large changes in the wing pitching moments and in aileron hinge moments. 

Another control issue regarding rolling moments is that ice accretions on the left and right wings 

are not typically symmetric. As a result of the ice asymmetry, flow separation in the spanwise 

direction could be different between the two sides of the wing, thus leading to considerable 

rolling moments. 

In reviewing the pressures distributions of the clean and iced wing configurations presented in 

figures 3-52 to 3-69, the following trends and features in the pressure data should be examined: 

• Location of the area centroid of the chordwise pressure distribution with respect to the 

25% MAC point of the wing. This determines pitching-moment behavior. 

3-75
• Change in pressure distribution in the spanwise direction. This affects spanwise load 

distribution and provides insight into spanwise flow separation. The spanwise load 

distribution also affects wing rolling moment. 

• The growth of separation bubbles with α. Sudden changes in bubble features are 

undesirable because they cause sudden changes in wing loads. 

• The shape of the pressure distribution over the upper and lower control surface. This is 

responsible for the observed behavior of control hinge moments. 

3.2.4.9 Clean Wing (Figures 3-52 to 3-54). 

• α = -4°: High suction was observed over the lower surface of the wing. Suction 

increased with spanwise distance and reached a maximum suction Cp of -3.53 at the 85% 

semispan station. The pressure data indicated that the flow was attached. 

• α = 0°: The flow remained attached. Due to the wing geometric twist, the average α was 

negative, resulting in negative lift. In all cases, the lower surface of the wing experienced 

suction that was lower near the root and higher near the tip. 

• α = 4°: For this α, αlocal was 4° at the root and 0° at the tip. The pressure over the wing 

upper surface was lower than over the lower surface, resulting in positive lift. Upper 

surface suction was higher near the wing root than near the wing tip. The flow remained 

attached. 

• α = 8°: High suction was observed over the wing upper surface at all spanwise stations. 

The suction Cp increased from -3.3 at 15% semispan to -3.5 at 50% semispan, and then it 

decreased to -2.7 at 85% semispan. The pressure data indicated that flow separation was 

not present over the wing for this α. 

• α = 12°: LE separation bubbles are evident for this near stall α. The extent of the 

separation bubbles increased from root to tip, while the suction over the bubble 

decreased. 

• α = 16°: A large LE separation bubble was observed near the wing root. The flat Cp 

curves corresponding to the wing upper surface at the 50% and 85% semispan stations 

are indicative of massive flow separation. The pressure data presented for α of 12° and 

16° show that flow separation (wing stall) was initiated at the tip of the wing and 

progressed toward the root. 

3.2.4.10 Iced Wing. 

Pressure distributions for the iced wing with the IRT castings and the simulated smooth and 

rough LEWICE ice shapes are discussed below. The experimental pressure distributions for the 

iced configurations tested exhibited leading-edge bubbles over the upper, lower, and in many 

cases, over both surfaces of the wing. As the α was increased, the bubble(s) grew longer in the 

3-76
downstream direction and eventually burst into massive flow separation as the streamlines failed 

to reattach to the wing surface. 

3.2.4.10.1 Icing Condition 1—Ice shapes: IRT-CS10, LS-CS10, LR-CS10 (Figures 3-52 to 3-54). 

• α = -4°: At 15% semispan, LE bubbles were evident over the wing lower surface for all 

ice shape cases. The suction over the bubble region was greater for the LEWICE ice 

shapes compared to the IRT ice shape casting. At the mid-semispan location, extensive 

flow separation was observed over the lower wing surface for the cases presented. The 

flow remained separated over the lower surface near the tip station, and the suction was 

reduced. An LE bubble was present over the lower surface for the LEWICE ice shapes. 

• α = 0°: Flow separation in the form of LE bubbles was present over both wing surfaces 

for all ice shapes tested. Suction over the lower surface was higher than over the upper 

surface. In most cases, the LEWICE ice shapes resulted in higher suction over the lower 

surface. The bubble extent was less near the wing root than at the outboard stations. 

Near the wing tip, flow separation extended to the wing TE. 

• α = 4°: For this angle of attack, suction was increased over the wing upper surface. The 

Cp on the lower surface varied from high negative near the LE, indicating the presence of 

separation bubbles to low negative and, in some cases, slightly positive near the TE. The 

extent of the upper surface bubbles decreased (separation was reduced) with spanwise 

distance from root to tip due to the negative wing twist, which lowered the αlocal over the 

outboard sections. 

• α = 8°: Long bubbles were observed over the inboard wing upper surface for all the ice 

shapes. The bubbles grew longer (chordwise) over the outboard sections of the wing and 

eventually, complete flow separation was observed near the wing tip. Suction over the 

bubbles for the LEWICE ice shapes was higher than obtained with the IRT ice shape 

casting. In most cases, the pressure distribution over the wing lower surface obtained 

with the LEWICE shapes exhibited bubble formation between 0% and approximately 

50% chord. 

• α = 12°: Large bubbles were observed at the 15% station, and complete flow separation 

was evident over the wing upper surface at the 50% and 85% semispan locations with all 

ice shapes tested. The IRT ice shape had lower suction over the two inboard stations than 

the LEWICE ice shapes. Bubbles were also present over the lower surface of the wing 

for the two LEWICE shapes. 

• α = 16°: At this angle of attack, extensive flow separation was observed over the upper 

surface of the wing at all spanwise stations. The only exception was the smooth 

LEWICE ice shapes for which the flow over the inboard (15% semispan) station was 

partially attached. 

3-77
3.2.4.10.2 Icing Condition 2—Ice Shapes: IRT-IS10, LS-IS10, LR-IS10 (Figures 3-55 to 3-57). 

• α = -4°: At 15% semispan, LE bubbles were evident over the wing lower surface for all 

ice shape cases. In addition, at the mid-semispan location, the bubbles grew longer. 

Furthermore, reduced suction and extensive flow separation were observed over the wing 

lower surface near the tip station. Note that the LE bubbles of the IRT-IS10 and LR-IS10 

cases were comparable at the 50% and 85% semispan stations, while the bubble of the 

LS-IS10 configuration produced higher suction. 

• α = 0°: Flow separation in the form of LE bubbles was present over both wing surfaces 

for all ice shapes tested. Suction over the lower surface was higher than over the upper 

surface. The IRT and LEWICE ice shapes resulted in comparable suction over the lower 

surface. The bubble extent was less near the wing root than at the outboard stations. 

Near the wing tip, flow separation extended to the wing TE. 

• α = 4°: For this angle of attack, suction was increased over the wing upper surface. The 

Cp on the lower surface varied from high negative near the LE, indicating the presence of 

separation bubbles, to low negative and, in some cases, slightly positive near the TE. 

Note that at the 15% semispan station, the suction on the wing upper surface, due to the 

IRT-IS10 ice shape, was lower than that produced by the LEWICE ice shapes. However, 

the difference in suction between the IRT and LEWICE ice shapes diminished with 

spanwise distance from root to tip. 

• α = 8°: Long bubbles were observed over the inboard wing upper surface for all ice 

shapes. The bubbles grew longer (chordwise) over the outboard sections of the wing and 

eventually complete flow separation was observed near the wing tip. Suction over the 

bubbles for the LEWICE ice shapes was higher than that obtained with the IRT ice shape 

casting near the wing root. In most cases, the pressure distribution over the wing lower 

surface obtained with the LEWICE ice shapes exhibited bubble formation between 0% 

and approximately 50% chord. 

• At the higher α of 12° and 16°, extensive flow separation was observed over the upper 

surface of the wing at all spanwise stations. The only exception was the flow at the 15% 

semispan station, which was partially attached for all ice shape configurations. 

3.2.4.10.3 Icing Condition 3—Ice Shapes: IRT-SC5, LS-SC5, LR-SC5 (Figures 3-58 to 3-60). 

• α = -4°: At 15% semispan, small LE bubbles were observed over the wing lower surface 

for all ice shape cases. The suction over the bubble region was comparable between the 

LEWICE and IRT ice shapes casting. The bubbles grew longer (chordwise) at the midsemispan station. At the 85% semispan location, extensive flow separation was observed 

over the wing lower surface for all cases presented. An LE bubble was present over the 

lower surface for all ice shapes. 

• α = 0°: Flow separation in the form of LE bubbles was present over both wing surfaces 

for all ice shapes tested. Suction over the lower surface was higher than over the upper 

3-78
surface. The bubble extent was less near the wing root than at the outboard stations. 

Near the wing tip, flow separation extended to the wing trailing edge. 

• For angles of attack of 4° and 8°, Cp distributions of all ice shape configurations were in 

good agreement with the clean case. At the two inboard locations, the minimum pressure 

near the LE for the iced wing was lower than the clean wing, indicating lower αe. The 

reduction in αe was due to the LE droop caused by the rime ice accretion. 

• α = 12°: At the 15% semispan station, only the IRT-SC5 ice shape produced a LE 

bubble on the wing upper surface. At the same time, observe that suction was increased 

for both LEWICE ice shapes. As the flow progressed to the 50% semispan location, the 

Cp distributions exhibited a bubble near the LE for all three iced wing configurations, 

where the suction over the bubble for the LS-SC5 case was higher than the other two SC5 

ice shapes. The LE bubbles grew in chordwise length from the 50% semispan station to 

the wing tip. 

• α = 16°: At this angle of attack, all ice shapes maintained partial attached flow over the 

wing suction surface at the two inboard stations. However, extensive flow separation 

was observed over the upper surface of the wing at the 85% location. 

3.2.4.10.4 Icing Condition 4—Ice Shapes: IRT-CS2, LS-CS2, LR-CS2 (Figures 3-61 to 3-63). 

• α = -4°: At 15% semispan, LE bubbles were evident over the wing lower surface for all 

ice shape cases. From root to mid-semispan stations, the bubble region extent increased. 

Extensive flow separation was observed over the wing lower surface near the tip station 

for all cases presented. 

• α = 0°: Flow separation in the form of LE bubbles was present over both wing surfaces 

for all ice shapes tested. Suction over the lower surface was higher than over the upper 

surface. Note that the bubble extent was less near the wing root than at the outboard 

stations. However, the LEWICE ice shapes produced separation bubbles of greater 

extent than the IRT ice shape. 

• α = 4°: For this angle of attack, suction was increased over the wing upper surface. 

Separation bubbles were observed over the lower surface between 0% and approximately 

20% chord. The extent of the upper surface bubbles decreased (separation was reduced) 

with spanwise distance from root to tip due to the negative wing twist, which lowered the 

αlocal over the outboard sections. Moreover, suction over the LE bubbles of the LEWICE 

ice shapes was higher than the IRT-CS2 case at all spanwise stations. 

• α = 8°: Once again, bubbles were observed over the inboard wing upper surface for all 

the ice shapes. The bubbles grew longer (chordwise) over the outboard sections of the 

wing. For the LEWICE ice shapes, suction over the bubbles was higher than obtained 

with the IRT ice shape casting at the 15% semispan station. However, at the midsemispan and near tip stations, the smooth and rough LEWICE ice shapes had lower 

suction over the separation bubbles than the IRT-CS2 ice shape. 

3-79
• α = 12°: Large LE bubbles were observed at the 15% semispan station, and complete 

flow separation was evident over the wing upper surface at the two outboard locations 

with all the ice shapes, except the IRT-CS2 ice shape for which the flow was partially 

attached near the wing tip. 

• α = 16°: At this angle of attack, extensive flow separation was observed over the upper 

surface of the wing at all spanwise stations. 

3.2.4.10.5 Icing Condition 5—Ice Shapes: IRT-CS22, LS-CS22N, LS-CS22S, LR-CS22N, and

LR-CS22S (Figures 3-64 to 3-66). 

• α = -4°: At 15% semispan, LE bubbles were evident over the wing lower surface for all 

ice shapes. Suction over the bubble region was greater for the LEWICE ice shapes 

compared to the IRT-CS22 case. Comparing the smooth and rough LEWICE ice shapes, 

it was observed that the addition of roughness to the ice horns produced bubbles with 

lower suction but of greater chordwise extent than obtained with the smooth LEWICE ice 

shape. At the mid-semispan location, extensive flow separation was observed over the 

wing lower surface for all cases, except the LS-CS22S and LR-CS22S ice shapes. The 

flow remained separated over the lower surface near the tip station. 

• α = 0°: Flow separation in the form of LE bubbles was present over both wing surfaces 

for all ice shapes tested. Suction over the lower surface was higher than over the upper 

surface. In most cases, the LEWICE ice shapes resulted in higher suction over the lower 

surface. Near the wing tip, flow was completely separated. 

• α = 4°: For this angle of attack, LE bubbles were present over both surfaces near the 

wing root. For the rough LEWICE ice shapes, suction over the bubbles was higher than 

their smooth counterparts. At the mid-semispan location, extensive flow separation was 

observed for the IRT and rough LEWICE ice shapes, while for the LS-CS22N and LSCS22S cases, the flow was partially attached. An LE bubble was evident for the IRTCS22 ice shape at the 85% semispan station. Flow remained separated near the wing tip 

for all LEWICE cases. 

• α = 8°: Long bubbles were observed over the inboard wing upper surface for all 

LEWICE ice shapes, whereas flow downstream of the IRT-CS22 horns was completely 

separated over the wing upper surface. For the LEWICE cases, the LE separation 

bubbles failed to reattach over the wing sections at the 50% and 85% semispan stations, 

and complete flow separation was observed. For the wing with LEWICE ice shapes, the 

pressure distribution over the lower surface exhibited bubble formation between 0% and 

approximately 50% chord. 

• α = 12°: Aft of IRT-CS22 upper ice horn, complete flow separation occurred. For all 

LEWICE ice shapes, large bubbles were observed at the 15% semispan station, and 

massive flow separation was evident over the wing upper surface at the 50% and 85% 

semispan locations. Note that the IRT ice shape had lower suction than the LEWICE ice 

shapes. Bubbles were also present over the lower surface of the wing for all ice shapes 

3-80
from root to tip. Yet, only the LS-CS22S ice shape maintained a lower surface bubble 

near the wing tip. 

• α = 16°: At this angle of attack, extensive flow separation was observed over the upper 

surface of the wing at all spanwise stations. Once again, all ice shapes produced an LE 

bubble over the wing pressure surface at the three spanwise stations, as shown in the 

figures. 

3.2.4.10.6 Icing Condition 6—Ice Shapes: IRT-IPSF22, LS-IPSF22, LR-IPSF22 (Figures 3-67

to 3-69). 

• α = -4°: At 15% semispan, LE bubbles were evident over the wing lower surface for all 

ice shape cases. The suction over the bubble region was greater for the LEWICE ice 

shapes compared to the IRT ice shape casting. At the mid-semispan location, bubbles of 

greater extent were observed over the wing lower surface for all cases presented. Over 

the lower surface, the flow was separated near the tip station for the IRT-IPSF22 

configuration, whereas an LE bubble was present for the LEWICE ice shapes. 

• α = 0°: Flow separation in the form of LE bubbles was present over both wing surfaces 

for all ice shapes tested. Suction over the lower surface was higher than that over the 

upper surface. In most cases, the LEWICE ice shapes resulted in higher suction over the 

lower surface. The bubble extent was less near the wing root than at the outboard 

stations. Near the wing tip, flow separation extended to the wing TE. 

• α = 4°: For this angle of attack, suction was increased over the wing upper surface. The 

Cp on the lower surface varied from high negative near the LE, indicating the presence of 

separation bubbles, to low negative and, in some cases, slightly positive near the TE. 

• α = 8°: Long bubbles were observed over the inboard wing upper surface for all ice 

shapes. The bubbles grew longer (chordwise) over the outboard sections of the wing, and 

eventually, complete flow separation was observed near the wing tip. Suction over the 

bubbles for the LEWICE ice shapes was higher than obtained with the IRT ice shape 

casting. Separation bubbles were also observed on the lower wing surface for all ice 

shapes. The lower surface bubble increased its chordwise length from root to tip. 

• α = 12°: Large bubbles were observed at the 15% station, and complete flow separation 

was evident over the wing upper surface at the 50% and 85% semispan locations for all 

ice shapes tested. The IRT ice shape had lower suction over the two inboard stations than 

the LEWICE ice shapes. Bubbles were also present over the lower surface of the wing 

for all ice shapes. 

• α = 16°: At this angle of attack, extensive flow separation was observed over the upper 

surface of the wing at all spanwise stations. The only exception was the smooth 

LEWICE ice shapes for which the flow over the inboard (15% semispan) station was 

partially attached. Once again, LE bubbles were observed for all ice shapes over the 

wing lower surface. 

3-81
3.2.5 Effect of Aileron Deflection. 

The effect of aileron deflection on the aerodynamic performance of the clean and iced wing 

is presented in figures 3-70 to 3-73 and in table 3-10. The main performance parameters 

presented include: 

• CL versus α for all aileron deflections (δA) 

• CH versus α for all δA

• CL versus δA for selected α

• CH versus δA for selected α

Another important parameter for assessing aileron performance is the change in rolling moment 

with δA and α. Rolling moment data were obtained during the experimental investigation at 

WSU. 

The experimental lift and hinge-moment data presented in figures 3-70 to 3-73 were obtained 

with a reflection plane wing model having a single aileron. Aircraft wings, however, have left 

and right ailerons that are not independent but are connected via cables to the control wheel. 

Thus, the control forces experienced by the pilot are due to the difference in the hinge moments 

generated by the left and right ailerons. To estimate the net aileron control force from the 

experimental data presented, the hinge moment for positive and negative aileron deflections must 

be combined. For example, using the results presented in figure 3-70(f) for the clean wing, α of 

0° and δA of 5°, the difference between CH for aileron deflection of -5° and CH for aileron 

deflection of 5° provides the net aileron control force. This procedure assumes that the up and 

down aileron deflection angles are the same. However, in some aircraft, left and right aileron 

deflection angles are not the same, and in such a case, the actual deflections for each aileron have 

to be used to compute the control force. Another assumption made regarding the 

superimposition of the experimental results is that ice accretions for the left and right wings are 

symmetric. In general, the left and right wing ice shapes will not be exactly the same, and this 

could cause different separation patterns over the left and right sides of the wing and the ailerons. 

3-82
(a) Lift coefficient (negative δA) (b) Lift coefficient (positive δA) 

(c) Hinge-moment coefficient (negative δA) (d) Hinge-moment coefficient (positive δA) 

(e) Lift coefficient vs δA (f) Hinge-moment coefficient vs δA

FIGURE 3-70. EFFECT OF AILERON DEFLECTION ON LIFT AND AILERON HINGEMOMENT COEFFICIENTS; CLEAN CONFIGURATION; Re = 1.8x106

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Lift coefficient

δA (deg)

0 (R006)

2.5 (R007)

5 (R008)

10 (R009)

15 (R010)

20 (R011)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.80

-0.60

-0.40

-0.20

0.00

0.20

0.40

0.60

0.80

1.00

Lift coefficient

δA (deg)

0 (R006)

-2.5 (R005)

-5 (R004)

-10 (R003)

-15 (R002)

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.60

-0.40

-0.20

0.00

0.20

0.40

0.60

0.80

1.00

1.20

1.40

1.60

1.80

Lift coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.20

-0.15

-0.10

-0.05

0.00

0.05

0.10

0.15

Hinge moment coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.15

-0.10

-0.05

0.00

0.05

0.10

0.15

0.20

Hinge Moment coefficient

δA (deg)

0 (R006)

-2.5 (R005)

-5 (R004)

-10 (R003)

-15 (R002)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.25

-0.20

-0.15

-0.10

-0.05

0.00

0.05

0.10

Hinge moment coefficient

δA (deg)

0 (R006)

2.5 (R007)

5 (R008)

10 (R009)

15 (R010)

20 (R011)

3-83
(a) Lift coefficient (negative δA) (b) Lift coefficient (positive δA) 

(c) Hinge-moment coefficient (negative δA) (d) Hinge-moment coefficient (positive δA) 

(e) Lift coefficient vs δA (f) Hinge-moment coefficient vs δA

FIGURE 3-71. EFFECT OF AILERON DEFLECTION ON LIFT AND AILERON HINGEMOMENT COEFFICIENTS; IRT-CS22 ICE SHAPE; Re = 1.8x106

-10 -5 0 5 10 15 20 2

Angle of attack (deg)

5

-0.30

-0.20

-0.10

0.00

0.10

0.20

0.30

0.40

0.50

0.60

0.70

Lift coefficient

δA (deg)

0 (R047)

2.5 (R042)

5 (R043)

10 (R044)

15 (R045)

20 (R046)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.30

-0.20

-0.10

0.00

0.10

0.20

0.30

0.40

0.50

0.60

Lift coefficient

δA (deg)

0 (R047)

-2.5 (R040)

-5 (R039)

-10 (R038)

-15 (R037)

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.40

-0.20

0.00

0.20

0.40

0.60

0.80

1.00

Lift coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.20

-0.15

-0.10

-0.05

0.00

0.05

0.10

Hinge moment coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.08

-0.06

-0.04

-0.02

0.00

0.02

0.04

0.06

0.08

0.10

Hinge moment coefficient

δA (deg)

0 (R047)

-2.5 (R040)

-5 (R039)

-10 (R038)

-15 (R037)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.20

-0.15

-0.10

-0.05

0.00

0.05

Hinge moment coefficient

δA (deg)

0 (R041)

2.5 (R042)

5 (R043)

10 (R044)

15 (R045)

20 (R046)

3-84
(a) Lift coefficient (negative δA) (b) Lift coefficient (positive δA) 

(c) Hinge-moment coefficient (negative δA) (d) Hinge-moment coefficient (positive δA) 

(e) Lift coefficient vs δA (f) Hinge-moment coefficient vs δA

FIGURE 3-72. EFFECT OF AILERON DEFLECTION ON LIFT AND AILERON HINGEMOMENT COEFFICIENTS; IRT-IPSF22 ICE SHAPE; Re = 1.8x106

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.60

-0.40

-0.20

0.00

0.20

0.40

0.60

0.80

Lift coefficient

δA (deg)

0 (R091)

-2.5 (R090)

-5 (R089)

-10 (R087)

-15 (R086)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.60

-0.40

-0.20

0.00

0.20

0.40

0.60

0.80

Lift coefficient

δA (deg)

0 (R091)

2.5 (R092)

5 (R093)

10 (R094)

15 (R095)

20 (R097)

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.60

-0.40

-0.20

0.00

0.20

0.40

0.60

0.80

1.00

1.20

Lift coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.20

-0.15

-0.10

-0.05

0.00

0.05

0.10

0.15

Hinge moment coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.10

-0.05

0.00

0.05

0.10

0.15

Hinge moment coefficient

δA (deg)

0 (R091)

-2.5 (R090)

-5 (R089)

-10 (R087)

-15 (R086)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.25

-0.20

-0.15

-0.10

-0.05

0.00

0.05

Hinge moment coefficient

δA (deg)

0 (R091)

2.5 (R092)

5 (R093)

10 (R094)

15 (R095)

20 (R097)

3-85
(a) Lift coefficient (negative δA) (b) Lift coefficient (positive δA) 

(c) Hinge-moment coefficient (negative δA) (d) Hinge-moment coefficient (positive δA) 

(e) Lift coefficient vs δA (f) Hinge-moment coefficient vs δA

FIGURE 3-73. EFFECT OF AILERON DEFLECTION ON LIFT AND AILERON HINGEMOMENT COEFFICIENTS; LR-IPSF22 ICE SHAPE; Re = 1.8x106

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

Lift coefficient

δA (deg)

0 (R374)

-2.5 (R370)

-5 (R371)

-10 (R372)

-15 (R373)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

Lift coefficient

δA (deg)

0 (R374)

2.5 (R369)

5 (R368)

10 (R367)

15 (R366)

20 (R365)

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.60

-0.40

-0.20

0.00

0.20

0.40

0.60

0.80

1.00

1.20

Lift coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.20

-0.15

-0.10

-0.05

0.00

0.05

0.10

0.15

Hinge moment coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.10

-0.05

0.00

0.05

0.10

0.15

Hinge moment coefficient

δA (deg)

0 (R374)

-2.5 (R370)

-5 (R371)

-10 (R372)

-15 (R373)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.25

-0.20

-0.15

-0.10

-0.05

0.00

0.05

Hinge moment coefficient

δA (deg)

0 (R374)

2.5 (R369)

5 (R368)

10 (R367)

15 (R366)

20 (R365)

3-86
TABLE 3-10. CL,stall OF ALL CONFIGURATIONS WITH 

AILERON DEFLECTION; ReMAC = 1.8x106 

δA 

Configuration -15° -10° -5° -2.5° 0° 2.5° 5° 10° 15° 20°

Clean 0.829 0.845 0.858 0.868 0.874 0.883 0.887 0.904 0.913 0.929 

IRT-CS10 0.491 0.506 0.522 0.526 0.536 0.544 0.533 0.573 0.593 0.613 

IRT-IS10 0.591 0.603 0.618 0.628 0.646 0.648 0.662 0.676 0.688 0.716 

IRT-SC5 0.862 0.874 0.888 0.893 0.896 0.908 0.918 0.930 0.948 0.961 

IRT-CS2 0.737 0.752 0.770 0.774 0.775 0.782 0.788 0.807 0.819 0.835 

IRT-CS22 0.005 0.016 0.025 0.028 0.033 0.036 0.037 0.039 0.039 0.037 

IRT-IPSF22 0.506 0.515 0.527 0.532 0.528 0.527 0.531 0.546 0.563 0.579 

LS-CS10 0.669 0.679 0.692 0.705 0.710 0.719 0.723 0.744 0.760 0.772 

LS-IS10 0.640 0.656 0.666 0.672 0.679 0.687 0.697 0.722 0.735 0.747 

LS-SC5 0.900 0.915 0.927 0.933 0.939 0.952 0.965 0.980 0.984 1.000 

LS-CS2 0.714 0.722 0.737 0.741 0.748 0.759 0.769 0.780 0.800 0.812 

LS-CS22N 0.450 0.462 0.476 0.476 0.484 0.490 0.496 0.505 0.512 0.520 

LS-CS22S 0.495 0.507 0.516 0.517 0.537 0.542 0.548 0.564 0.575 0.601 

LS-IPSF22 0.635 0.648 0.660 0.667 0.676 0.683 0.697 0.713 0.737 0.746 

LR-CS10 0.564 0.584 0.597 0.607 0.613 0.624 0.635 0.647 0.669 0.683 

LR-IS10 0.589 0.605 0.627 0.636 0.644 0.651 0.661 0.678 0.694 0.711 

LR-SC5 0.921 0.939 0.955 0.962 0.980 0.982 0.994 1.012 1.027 1.044 

LR-CS2 0.677 0.691 0.703 0.709 0.717 0.730 0.736 0.754 0.770 0.783 

LR-CS22N 0.406 0.405 0.409 0.406 0.407 0.412 0.413 0.416 0.424 0.427 

LR-CS22S 0.352 0.351 0.349 0.346 0.342 0.344 0.343 0.352 0.351 0.356 

LR-IPSF22 0.566 0.577 0.590 0.602 0.612 0.617 0.626 0.646 0.659 0.680 

3.2.5.1 General Comments Regarding Aileron Controls. 

Ailerons are designed to meet aircraft roll performance criteria in terms of roll rate (typically 10 

to 15 degrees per second) and attitude. In addition, the ailerons should be able to provide 

sufficient roll control to handle special aircraft situations such as an engine-out case or steady 

heading sideslip. Aileron forces should remain within the capabilities of the pilot as defined by 

the certification requirements [11], which state that the maximum permissible aileron control 

force is 50 lb with two hands on the control wheel and 25 lb for a single hand on the control 

wheel (autopilot settings vary with design but often a 20 lb limit is used). 

The change in aileron hinge moments with δA and α has a direct impact on aileron control forces. 

The slope of CH with δA, CH,δA, also known as the control heaviness parameter, is of primary 

importance in evaluating control surface behavior and control forces. In general, aileron designs 

have negative CH,δA so that the aileron has the tendency to return to its undeflected position. The 

hinge-moment slope, CH,α, also known as the control floating parameter, affects the change in the 

control force resulting from the response of the aircraft to the control movement. When the 

ailerons are deflected and wing roll is initiated, the local α of the upward and downward moving 

wings is changed (typically by a couple of degrees for small deflection), and the control force 

required to maintain a steady maneuver is either greater or less than the control force required to 

3-87
initiate the maneuver, depending on the sign and magnitude of CH,α. Control surface designers 

usually aim at small positive or negative values for CH,α to avoid large differences in control 

forces during maneuver [18 and 19]. The effect of CH,α on control forces is of secondary 

importance, particularly for small aileron deflection and roll rates that cause small changes in α. 

For large deflections, however, which can produce large roll rates and therefore a large change in 

α, CH,α could have a notable effect on control forces. 

Once the clean wing aileron design requirements have been satisfied, the effect of ice shapes on 

aileron performance should be assessed for unwanted changes in control behavior. In general, 

the presence of ice on the wing leading edge can result in early flow separation over the wing 

and control surfaces, and the iced wing lift with δA could change with respect to the clean wing. 

It is important that the lift of the iced wing exhibits the expected behavior with positive and 

negative aileron deflections, although some loss in aileron effectiveness is expected. 

The effect of ice shapes on aileron hinge moments (i.e., control forces) and on the overall 

behavior of the hinge-moment curve should be evaluated. With iced wings, the growth in hinge 

moment with α (region B as defined in figures 3-26(a) and 3-26(b)) takes place at a lower α with 

respect to the clean wing, as shown in figures 3-26(c) and 3-26(d). For such a case, a large 

amplitude roll maneuver could change the α sufficiently to place the aileron on the down moving 

side of the wing in the steep region of the hinge-moment curve. This could cause a considerable 

increase in control force. If, in addition to the increase in control force, the aileron effectiveness 

is reduced because of flow separation over the ailerons caused by the ice shape, then a situation 

could develop where recovery may be difficult. 

In assessing aileron performance for iced wings, the magnitude of CH and the slope of CH with 

respect to α and δA should be considered. Ideally, the magnitude of CH should be balanced; 

small enough to avoid excessively high forces, yet large enough to avoid very small forces. The 

actual permissible values of CH will be a function of the aerodynamic and mechanical design of 

the control, as well as certification requirements. Any change in the slope of CH with α or δA

should also be small and gradual. 

3.2.5.2 Clean Wing. 

For a reflection place model, positive aileron deflections increase wing lift and reduce the α,stall, 

while negative deflections have the opposite effect. Note that large positive aileron deflections 

at high angles of attack should be evaluated carefully since they increase the angle of attack over 

the outboard portion of the wing and can result in early tip stall. 

For the clean wing tested, the aileron remained effective for all α and aileron deflections, i.e., 

positive aileron deflection (TE down) increased wing lift, while negative deflection resulted in 

lift reduction, as expected. In addition, as the aileron deflection was increased, the lift increment 

(for aileron TE down) or decrement (for aileron TE up) increased, as shown in figures 3-70(a) 

and 3-70(b). 

The aileron hinge moments presented in figures 3-70(c) and 3-70(d) exhibited gradual change 

with α for all aileron deflections tested. The slope, CH,α, for all three regions A, B, and C (as 

3-88
defined in figure 3-26) of the CH curve were negative, indicating that the hinge moment became 

less positive or more negative as α was increased from -8° to 20°. On a number of cases, a 

change in the sign of CH was observed in the hinge-moment data. Consider, for example, the 

case in figure 3-70(c), corresponding to δA of -15° (TE up). For this case, a change in α from 

10° to 16° resulted in a change in aileron hinge moments from 0.07 to approximately -0.03. The 

change in sign of CH occurred near wing stall, corresponding to α of approximately 13.5°. For α

between 10° and 13.5° and for δA of -15°, the hinge moment was positive, indicating that the 

aileron trailing edge had the tendency to return to its undeflected position. However, at α greater 

than 13.5°, the aileron had the tendency to maintain its deflected position due to extensive flow 

separation over the upper surface of the aileron. Flow conditions which force a control surface 

to maintain its deflection status result in control force reversal. However, with aileron controls, 

such a condition over one of the ailerons may not be a problem if the other aileron (the one 

deflected TE down in this case) remains effective. Considering the 15° aileron deflection case in 

figure 3-70(d), it is observed that for α of 13.5°, the hinge moment for the TE down aileron was 

approximately -0.18, i.e., the TE down aileron had a strong tendency to return to its neutral 

position. Thus, the combined hinge moment and, therefore, the net control force due to both 

ailerons maintained the correct behavior (i.e., the controls had the tendency to return to the 

neutral position). 

The change in CL with δA depicted in figure 3-70(e) was practically linear for α in the range of -

8° to 16°. The change in lift as δA was increased from -15° to 20°, while maintaining a constant 

α, was in the range of 0.1 to 0.2, depending on α. The higher lift increments were obtained for α

in the linear lift range where the flow over the control surface was mostly attached. 

The change in control hinge moment with δA is demonstrated in figure 3-70(f). The aileron CH

was a function of the prevailing flow conditions over the aileron. Consider the line in 

figure 3-70(f) corresponding to α of 0°. At this α, the flow was attached on both surfaces of the 

wing for all aileron deflections. The results show that for negative δA (aileron TE up), CH was 

positive, that is the aileron had the tendency to return to its undeflected position. For positive 

aileron deflections (TE down), CH is negative, which once again indicates that the aileron had the 

tendency to return to the neutral position. The linear behavior of the curve shows that as δA was 

increased, the moment and, therefore, the control force were increased in a linear fashion. Note 

that for a full-span wing, the net aileron control force can be obtained from the results presented 

in figure 3-70(f) by computing the change in CH, corresponding to the same positive and negative 

aileron angular deflections. For example, at α of 0°, the change in the magnitude of CH

corresponding to aileron deflections of ±5°, was approximately 0.075; CH for TE up aileron was 

0.032, while CH for TE down aileron was -0.043. 

Now consider the CH versus δA curve corresponding to α of -8°. At this α, the upper surface of 

the wing was the pressure side, while the lower surface was the suction side. For negative 

aileron deflections, the hinge moment was positive as the aileron upper surface experienced 

higher pressure than the lower surface. For small positive aileron deflections, the pressure over 

the aileron upper surface remained higher than over lower surface, which now had lower suction, 

and the hinge moment remained positive. As the aileron approached maximum positive 

deflection, the pressure over the upper surface did not change significantly but the suction over 

3-89
the lower surface was considerably reduced. Thus, the magnitude of the positive hinge moment 

was reduced. The reason for the nonlinearity in CH versus δA curve was mild flow separation 

over the aileron lower surface. 

The results presented in figure 3-70(f) demonstrate approximately linear change in CH with δA

for α corresponding to the linear lift range. For these α, the change in aileron hinge moment was 

approximately 0.21 as δA was increased from -15° to 15°. At high negative or positive α, the 

change in hinge moment with aileron deflection was not linear. This is demonstrated in figure 

3.70(f) for α of 16°, 12°, and -8°. 

3.2.5.3 IRT-CS22 Ice Shape. 

The effects of the 22.5-min glaze ice accretion with the large scallop features on the aerodynamic 

performance of the swept wing for all aileron deflections tested is presented in figure 3-71. The 

experimental results obtained indicate the following: 

• Lift increase or decrease with δA was as expected for most α as demonstrated in 

figures 3-71(a) and 3-71(b). For α-range of 1° to 6°, however, negative aileron 

deflections increased lift while positive deflections had practically no effect on lift. This 

was mainly due to massive flow separation over the wing and control surfaces caused by 

the large size (2- to 3-in.) ice horns. 

• The maximum positive CH was 0.1 and occurred for δA of -15° and α of -8° as shown in 

figure 3-71(c). The maximum negative hinge-moment coefficient was -0.19 at α of 20o 

and δA of 20o as evident from figure 3-71(d). Corresponding maximum and minimum 

values for the clean wing were 0.165 and -0.22. Thus, the iced wing hinge moments were 

within the maximum and minimum limits of the clean wing hinge moments. It is 

important, however, to point out that for some aircraft, the maximum hinge moment for 

the clean wing are typically not reached either due to the aircraft natural angle of attack 

limits or due to an artificial stall barrier. In such a case, the iced wing hinge moments 

should be compared to the actual clean wing hinge moment limits. 

• A change in the sign of CH was observed for all negative aileron deflections and for 

α-range of 2° to 5°, as shown in figure 3-71(c). For positive aileron deflections in the 

range of 0° to 5°, the change in the sign of CH occurred between α of 1° and 3°, as 

demonstrated in figure 3-71(d). For larger positive aileron deflections, the sign of CH

was reversed twice near α of 0°. As discussed previously, for the clean wing case, a sign 

change in CH indicates that the aileron has the tendency to maintain its deflected position. 

To determine the net aileron control hinge moment and its direction for the combined left 

and right aileron system, the CH for both positive and negative aileron deflections should 

be examined. 

• As demonstrated in figure 3-71(e), the change in wing lift with δA for constant α was 

practically linear for all α presented. The lines had a small positive slope for most α. 

3-90
However, for α-range of 0° to 4°, the slope was nearly zero and, in some cases, slightly 

negative, indicating that the aileron was not effective in changing lift. 

• The behavior of CH with δA was notably nonlinear for all α presented in figure 3-71(f). 

Given an α and assuming attached flow, CH will be positive for negative aileron 

deflections (TE up) since the aileron will tend to move to its undeflected position. For 

positive aileron deflections, the sign of CH will eventually become negative. For fixed 

positive aileron deflection, the pressure over the aileron lower surface and the suction 

over the upper surface should increase as the α is increased from -8° to 0° (note that such 

a change in α with fixed δA is not practical under flight conditions). The CH should 

become progressively more negative and the hinge-moment curves will move further 

apart. This was not the case with the IRT-CS22 ice shape, as shown in figure 3-71(f) for 

α-range of -8° to 0°. The intersection in the hinge-moment curves corresponding to these 

α near δA of 5° is indicative of extensive flow separation over the control surface. 

3.2.5.4 IRT-IPSF22 and LR-IPSF22 Ice Shapes. 

Aerodynamic performance with aileron deflection for the 22.5-min glaze ice shape casting IRTIPSF22 and the corresponding rough LEWICE ice shape LR-IPSF22 are presented in 

figures 3-72 and 3-73 respectively. These ice accretions were representative of ice protection 

system failure cases for the wing model tested. The results presented in figures 3-72 and 3-73 

indicate the following: 

• Increase or decrease in lift with δA was as expected for all aileron deflections and α, as 

demonstrated in figures 3-72(a), 3-72(b), 3-73(a), and 3-73(b). 

• The maximum positive CH for both ice shapes was approximately 0.14 and occurred for 

δA of -15° and α of -8°, while the maximum negative CH was -0.2 at α of 20° and δA of 

20°, as shown in figures 3-72(c), 3-72(d), 3-73(c), and 3-73(d). Once again, the iced 

wing hinge moments were within the maximum and minimum hinge-moment bounds of 

the clean wing. 

• The change in lift with δA for the IRT-IPSF22 and LR-IPSF22 ice shapes was practically 

linear for all α as shown in figures 3-72(e) and 3-73(e). For all cases presented in these 

figures, the CL versus δA curves had a small positive slope indicating a monotonic 

increase in lift with aileron deflection. 

• For most of the cases presented in figures 3-72(f) and 3-73(f), the change in the hingemoment coefficient with aileron deflection exhibited nonlinear behavior mainly due to 

flow separation over the aileron. 

3.2.6 Effect of Large Scallop Features. 

In general, ice accretions on swept wings can be classified as complete scallops, incomplete 

scallops, and no-scallop, as discussed in references 2, 4, 20, and 21. Schematics of these three 

types of ice accretions are provided in figure 3-74. 

3-91
Complete scallops are glaze ice accretions that appear only on swept wings and are characterized 

by the presence of scallop tips extending from the attachment line. The scallop tips have a 

particular shape, height, and spacing. Incomplete scallops are glaze ice shapes that also appear 

only on swept wings and where scallops tips form beginning at some distance from the 

attachment line. No-scallops are ice accretions where scallop tips are not present. 

Scallop tips are made of glaze ice feathers, which develop from roughness elements that form 

during the ice accretion process as shown in figure 3-75(a). The feathers have a preferred 

direction of growth that is perpendicular to the external streamlines, as shown in figures 3-75(b) 

and 3-75(c). Ice feathers join along the preferred direction to form ridges, with incipient scallop 

tips formed by the feathers at the end of each ridge. As the ridges grow, they form scallop tips. 

As the ice accretion grows, the scallop tips merge with adjacent scallop tips by joining at the top. 

This increases both the size of the scallop tips and the gap between them. Detailed descriptions 

and experimental data for scallop ice accretions can be found in references 2, 4, 20, and 21. 

The type of ice accretion that will form on a swept wing depends on the airfoil geometry, sweep 

angle, flow, and tunnel conditions. Icing conditions that lead to scalloped ice features can 

produce ice shapes with complex surface characteristics such as the IRT-CS10 and IRT-CS22 

cases shown in figure 2-33. 

Ice shapes with large scallop formations contain gaps between scallops, as demonstrated in 

figure 2-33(b) for the IRT-CS22 case. Aerodynamicists have debated over the years the effects 

of scallop features on aerodynamic performance. This is of practical interest because state-ofthe-art ice accretion codes are not able to produce scalloped ice shapes. Simulated ice shapes for 

swept wings developed with the use of ice accretion codes or other empirical means have solid 

horns, as shown in figure 2-34. If the scallop-gap features in the ice horns have a significant 

impact on aerodynamic performance, then methods should be explored for incorporating their 

effects in simulated ice shapes. 

To investigate the effect of scalloped ice shapes on aerodynamic performance, a limited study 

was conducted with the IRT-CS22 ice shape. In this study, the gaps between scallops were 

progressively filled with a modeling compound to produce an ice shape with solid horns, as 

shown in figure 3-76. To reduce the aerodynamic loads on the modeling compound, the 

experimental study with the baseline and modified IRT-CS22 ice shapes was conducted at a 

lower airspeed for which the Reynolds number based on MAC was 1.0 million. Experimental 

results from these tests are presented in figure 3-77. The first set of tests was conducted with the 

middle 25-in segment of the IRT-CS22 ice shape filled with the modeling compound (case 1, 

R516 in figure 3-77). Next, the gaps in the 18-in tip segment were filled (case 2, R517 in 

figure 3-77), and finally, the gaps in all three ice segments were filled (case 3, R518 in figure 3-

77). Each time a new set of force, moment, and pressure data was obtained. 

3-92
(a) Complete scallop (a) Attachment line zone, glaze ice feathers zone 

and critical distance (view from direction 

normal to leading edge) 

(b) Incomplete scallop (b) Attachment line zone, glaze ice feathers zone 

and critical distance (overall view) 

(c) No scallops (c) Feather detail 

FIGURE 3-74. ICE ACCRETION ON 

A SWEPT WING AT GLAZE ICE 

CONDITIONS [21] 

FIGURE 3-75. DEVELOPMENT OF 

ROUGHNESS ELEMENTS AND ICE 

FEATHERS DURING ICE ACCRETION 

ON A SWEPT WING [20 AND 21] 

Glaze Ice 

Feather 

Zone 

Attachment Line 

Proper 

dcr 

Roughness 

Elements 

Attachment 

Line Zone 

Complete Scallop

Attachment Line Proper 

Air Flow

Attachment 

Line Proper 

dcr critical 

distance 

Preferred Direction of Growth 

Streamline

Air Flow 

Inclination 

into the 

Flow 

Λ

V∞

Streamline 

Glaze Ice Feathers 

Forming 

Λ

Sweep Angle 

Scallop Tips 

No Scallop Tips 

3-93
(a) Filling in the tip segment (d) Filling in all segments – Front View 

(b) Filling in the middle and tip segments (e) Filling in all segments – Side View 

(Pressure) 

(c) Filling in the root segment (f) Filling in all segments – Side View 

(Suction) 

FIGURE 3-76. ICING RESEARCH TUNNEL CS22 ICE SHAPE FILLED WITH 

MODELING COMPOUND 

3-94
(a) Lift coefficient (d) Cp vs. x/c, 15%-semispan, α = 5°

(b) Drag coefficient (e) Cp vs. x/c, 50%-semispan, α = 5°

(c) Hinge-moment coefficient (f) Cp vs. x/c, 85%-semispan, α = 5°

FIGURE 3-77. EFFECT OF SCALLOP FEATURES ON AERODYNAMIC 

PERFORMANCE; IRT-CS22 ICE SHAPE WITH AND WITHOUT MODELING 

COMPOUND; Re = 1.8x106

; δA = 0°

0 5 10 15 20 25

Angle of attack (deg)

0.0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

Lift coefficient

No Filling (R113)

Mid Section Only (R516)

Mid and Tip Sections (R517)

All Sections (R518)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

0.1

0.0

-0.1

-0.2

-0.3

-0.4

-0.5

-0.6

-0.7

-0.8

-0.9

-1.0

-1.1

Pressure coefficient

Baseline IRT-CS22 (R113)

Root, Mid and Tip Section Gaps Closed (R518)

0 5 10 15 20 25

Angle of attack (deg)

0.20

0.25

0.30

0.35

0.40

0.45

Drag coefficient

No Filling (R113)

Mid Section Only (R516)

Mid and Tip Sections (R517)

All Sections (R518)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

0.1

0.0

-0.1

-0.2

-0.3

-0.4

-0.5

-0.6

-0.7

-0.8

-0.9

-1.0

-1.1

Pressure coefficient

Baseline IRT-CS22 (R113)

Root, Mid and Tip Section Gaps Closed (R518)

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

0.1

0.0

-0.1

-0.2

-0.3

-0.4

-0.5

-0.6

-0.7

-0.8

-0.9

-1.0

-1.1

Pressure coefficient

Baseline IRT-CS22 (R113)

Root, Mid and Tip Section Gaps Closed (R518)

0 5 10 15 20 25

Angle of attack (deg)

-0.08

-0.06

-0.04

-0.02

0.00

0.02

Hinge moment coefficient

No Filling (R113)

Mid Section Only (R516)

Mid and Tip Sections (R517)

All Sections (R518)

3-95
Lift, drag, hinge-moment, and pressure coefficients for the baseline IRT-CS22 ice shape (R113 

in figure 3-77) and for the three cases where the gaps between scallops were partially or 

completely filled with modeling compound are compared in figure 3-77. The results presented 

in this figure show a progressive increase in CL for α-range of -2° to 9° as more gaps in the ice 

horns were filled with the modeling compound. The maximum gain in CL was obtained for 

case 3 and ranged from 0.022 to 0.073 with respect to that of the baseline IRT-CS22 case. The 

change in CD with respect to the baseline ice shape was small but measurable for all three cases 

tested. For case 1, the change in the CH with respect to the baseline IRT-CS22 ice shape was 

small. However, for cases 2 and 3, the hinge moments were less negative (i.e., the aileron 

trailing edge had less of a tendency to move up). 

The exact mechanism for the observed aerodynamic effects caused by the gaps between the ice 

scallops is not known. However, the pressure data at 15%, 50%, and 85% semispan stations for 

case 3 (R518) presented in figure 3-77 for α of 5° offer some clues. The pressure distributions 

for the case where all the gaps between scallops were filled with the modeling compound 

exhibited increased suction on both the upper and lower wing surfaces with respect to the 

baseline IRT-CS22 ice shape. 

For large glaze ice accretions, the region between the horns is typically a high-pressure region 

because the air flow slows down within the cavity formed by the horns. The regions downstream 

of the upper and lower horns are in general low-pressure regions with small or large separation 

bubbles. With scalloped ice shapes, the gaps between the scallops allow the high pressure 

between the horns to leak to the low-pressure region downstream of the horns, thus reducing the 

suction near the wing LE. This is clearly evident in the 15% and 50% semispan pressure 

distributions presented in figures 3-77(d) and 3-77(e). At the 80% semispan station, the IRTCS22 ice shape with modeling compound produced lower suction peaks near the wing LE than 

the baseline ice shape. This was mainly due to flow separation that was more extensive for the 

case of the modified ice shape. Since CL depends on the pressure difference between the upper 

and lower wing surfaces, as this difference was reduced due to flow leakage in the case of the 

baseline IRT-CS22 ice shape, the net lift was also decreased. 

The observed lift increase with the modified IRT-CS22 ice shape (case 3, R518) in figure 3-77(a) 

can be explained by direct examination of the pressure distributions presented in figures 3-77(d) 

to 3-77(f). The pressure distributions show that, for the 15% semispan station, case 3 produced 

72% more positive lift than the baseline case. At the 50% semispan station, case 3 resulted in 

18% less negative lift than the baseline ice shape. At the 85% semispan location, both 

configurations resulted in about the same amount of negative lift. Thus, for case 3 where all 

gaps between the scallops were closed, the increase in suction over the wing resulted in a net 

increase in CL with respect to the baseline ice shape. 

In summary, the limited study conducted with the glaze IRT-CS22 scalloped ice shape showed 

that the scallop features and in particular the gaps between scallops can result in greater loss of 

lift than ice shapes with solid horns. It must be stressed, however, that these results should not 

be generalized until further studies are conducted with more scalloped type ice shapes. 

Furthermore, the results presented here may only be applicable to ice shapes with large scalloped 

features and for ice shapes with small scallops, the effects may not be as significant. 

3-96
3.2.7 Reynolds Number Effects. 

Reynolds number effects on the clean and selected iced wing configurations is presented in 

figure 3-78. Reynolds numbers included 0.5, 1.0, 1.5, and 1.8 million were computed based on 

the wing mean aerodynamic chord of 1.56 ft. Corresponding tunnel airspeeds were 42, 84, 128, 

and 152 mph. 

For the clean wing, experimental results for Reynolds numbers in the range of 0.5 to 1.8 million 

are presented in figure 3-78. For the low Reynolds number of 0.5 million, the CL,stall was 

approximately 0.87 and αstall was approximately 16°. For higher Reynolds numbers of 1.0 and 

1.5 million, both CL,stall and αstall decreased and were approximately 0.84 and 15° respectively. 

However, as the Reynolds number was increased to 1.8 million, CL,stall increased to 0.87 while 

αstall decreased to 13.8°. 

Reynolds number effects on CL of the IRT-IPSF22, LS-IPSF22, LR-IPSF22, IRT-SC5, and LRSC5 configurations are presented in figures 3-78b to 3-78(f). In general, the Reynolds number 

effects on CL of the IRT-IPSF22 ice shape was small, as demonstrated in figure 3-78(b). The 

effects of Reynolds number on the near stall lift performance of the LS-IPSF22 and LR-IPSF22 

ice shapes were more pronounced than for the IRT-IPSF22 case, as demonstrated in figures 3-

78(c) and 3-78(d). For the LR-IPSF22 configuration, as the Reynolds number was increased 

from 0.5 to 1.8 million, CL,stall decreased while αstall remained approximately the same. 

For the 5-min rime IRT-SC5 and LR-SC5 ice shapes, the effects of Reynolds number on CL are 

demonstrated in figures 3-78(e) and 3-78(f) respectively. Most of the change in lift with 

Reynolds number was observed in the near-stall and poststall flow regimes. In general, CL,stall

decreased as the Reynolds number was increased from 1.0 to 1.8 million. The behavior of CL 

for the low Reynolds number case of 0.5 million was similar to that for the 1.5 million Reynolds 

number. The Reynolds number effects observed with the 5-min rime ice cases were mainly due 

to the change in the location of flow separation over the rough ice shapes. At low Reynolds 

numbers, roughness can enhance or reduce performance, depending on the value of Reynolds 

number, by changing the state of the viscous boundary layer. At flight Reynolds number, 

however, even small levels of roughness can result in considerable performance losses, as was 

demonstrated in reference 18. The Reynolds number range explored in this study was small and 

considerably lower than that experienced by full-scale wings at flight conditions. It is not known 

how much different flight test results would be. 

3-97
 

(a) Lift coefficient; Clean configuration (b) Lift coefficient; IRT-IPSF22 ice shape 

(c) Lift coefficient; LS-IPSF22 ice shape (d) Lift coefficient; LR-IPSF22 ice shape 

(e) Lift coefficient; IRT-SC5 ice shape (f) Lift coefficient; LR-SC5 ice shape 

FIGURE 3-78. REYNOLDS NUMBER EFFECT ON CL; CLEAN AND ICED WING; δA = 0°

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Lift coefficient

Re=0.5x106 (R102)

Re=1x106 (R101)

Re=1.5x106 (R100)

Re=1.8x106 (R006)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

Lift coefficient

Re=0.5x106 (R567)

Re=1x106 (R566)

Re=1.5x106 (R565)

Re=1.8x106 (R091)

-10 -5 0 5 10 15 20 2

Angle of attack (deg)

5

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

Lift coefficient

Re=0.5x106 (R571)

Re=1x106 (R570)

Re=1.5x106 (R569)

Re=1.8x106 (R348)

-10 -5 0 5 10 15 20 2

Angle of attack (deg)

5

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

Lift coefficient

Re=0.5x106 (R423)

Re=1x106 (R422)

Re=1.5x106 (R421)

Re=1.8x106 (R374)

-10 -5 0 5 10 15 20 2

Angle of attack (deg)

-10 -5 0 5 10 15 20 2

Angle of attack (deg)

5

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Lift coefficient

Re=0.5x106 (R119)

Re=1x106 (R118)

Re=1.5x106 (R117)

Re=1.8x106 (R055)

5

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

1.2

Lift coefficient

Re=0.5x106 (R429)

Re=1x106 (R428)

Re=1.5x106 (R427)

Re=1.8x106 (R216)

3-98
4. CONCLUSIONS. 

Wind tunnel tests were conducted to investigate the effect of 20 ice shapes on the aerodynamic 

performance of a swept finite wing model. The wing consisted of an 8.7% thick airfoil section, 

which remained constant from root to tip. The wing was tapered and had a leading-edge (LE) 

sweep of 28°, a trailing-edge sweep of 15.6°, aspect ratio of 6.8, and -4° geometric twist at the 

tip (washout). Tests were performed with the clean wing, six ice shapes castings obtained from 

ice accretion experiments at the NASA Glenn Icing Research Tunnel (IRT), and seven smooth 

and seven rough LEWICE ice shapes. The LEWICE ice shapes were obtained for the same icing 

conditions as those used in the IRT ice accretion tests. One LEWICE ice shape was defined for 

each icing condition using airfoils sections normal to the wing LE at five spanwise stations. For 

icing condition 5, an additional LEWICE ice shape was defined using streamwise airfoil 

sections, as discussed in section 2.2.3 of this report. Roughness effects for the LEWICE ice 

shapes were simulated with 36-size grit. Tests were performed with the clean and iced wing in 

the low-speed 7- x 10-ft wind tunnel at Wichita State University. A complete set of force and 

moment coefficients were obtained along with aileron hinge moments and pressure distributions 

for a range of test conditions. Test conditions included Reynolds number of 1.8 million based on 

the wing mean aerodynamic chord yielding test Reynolds numbers well below typical flight 

Reynolds numbers, angles of attack in the range of -8° to 20°, and aileron deflections of -15°, 

-10°, -5°, -2.5°, 0°, 2.5°, 5°, 10°, 15°, and 20°. From the results presented, the following 

conclusions are drawn. 

4.1 PERFORMANCE OF CLEAN WING. 

The clean wing CL,stall was 0.87 and occurred at α of 13.8°. The linear lift slope was 0.077 per 

degree or 4.41 per radian. Drag coefficient (CD) varied from a minimum value of 0.0063 at α of 

1° to 0.15 at α of 13.8°. The maximum lift to drag ratio for the swept wing was 27.5 at α of 3.2°. 

The hinge-moment coefficient (CH) with the aileron in the neutral position varied from 0.079 at 

α of -8° to -0.02 at α of 0°, to -0.098 at α of 15.8°. The CH increased gradually through the 

nonlinear portion of the lift curve and attained a maximum slope of -0.0195 per degree. The 

clean wing experienced a leading edge long bubble/vortex stall, which progressed from the 

center of the wing to the tip and then to the root as α was increased. 

4.2 EFFECTS OF IRT ICE SHAPE CASTINGS. 

The six IRT ice shape castings included 2-, 10- and 22.5-min glaze ice accretions with 

incomplete and complete scallop features and a 5-min rime ice shape. The glaze ice shapes 

resulted in 11% to 93.6% reduction in CL,stall and 8% to nearly 56.5% reduction in αstall with 

respect to the clean wing. These ice shapes increased the clean wing CD,min by 200% to 3533%. 

In many cases, the behavior of the aileron CH for the clean wing was considerably altered by the 

glaze ice accretions. However, in all cases, CH of the iced wing remained within the maximum 

and minimum limits defined by the clean wing aileron CH. The 5-min rime ice shape improved 

the wing CL,stall by approximately 3% and the αstall by 14.5%. The iced wing CD,min was 133% 

greater than that of the clean wing. Aileron CH for the 5-min rime case were similar to that of the 

clean wing. 

4-1
For the IRT-CS22 ice shape with the large scallops, the results presented showed that the gaps 

between the scallops increased lift degradation. This was due to flow leakage from the near 

stagnation region between the ice shape horns to the low-pressure region downstream of the ice 

horns. This is a preliminary finding and requires further investigation before more general 

conclusions can be drawn regarding the effects of gaps in scalloped ice shapes. 

4.3 DEFINITION OF LEWICE ICE SHAPES. 

A methodology for defining three-dimensional simulated ice shapes for a swept finite wing using 

the two-dimensional LEWICE ice accretion code was presented. 

4.4 EFFECTS OF SMOOTH LEWICE ICE SHAPES. 

The glaze ice shapes resulted in 14% to 44% reduction in CL,stall. The αstall for the iced wing 

ranged from 13.4° to 14.8°. The iced wing CD,min was 183% to 2367% greater than that of the 

clean wing. In most cases, the aileron CH for the wing with the glaze LEWICE ice shapes were 

more positive or more negative than that of the clean wing for α-range of -7° to +13°. However, 

in all cases, the CH of the iced wing remained within the maximum and minimum limits defined 

by the clean wing aileron CH. The 5-min rime LEWICE ice shape improved CL,stall by 

approximately 8% and the clean wing αstall was increased by 1°. The CD,min for this ice shape 

was 67% greater than that of the clean wing. Aileron CH for the 5-min rime LEWICE ice shape 

were similar to that of the clean wing. 

4.5 EFFECTS OF ROUGH LEWICE ICE SHAPES. 

The rough glaze ice shapes resulted in 17% to 61% reduction in CL,stall. The αstall for the iced 

wing ranged from 12.6° to 13.7°. The iced wing CD,min was 233% to 3317% greater than that of 

the clean wing. In most cases, the aileron CH for the wing with the rough glaze LEWICE ice 

shapes were more positive or more negative than those of clean wing for α-range of -7° to +13°. 

However, in all cases, CH of the iced wing remained within the maximum and minimum limits 

defined by the clean wing aileron CH. The 5-min rime LEWICE ice shape improved the clean 

wing CL,stall by approximately 13% and increase αstall by 2°. The CD,min for this ice shape was 

117% greater than that of the clean wing. Aileron CH for the 5-min rime LEWICE ice shape 

were in general similar to that of the clean wing. 

4.6 ROUGH VERSUS SMOOTH LEWICE ICE SHAPES. 

In general, the smooth and rough LEWICE ice shapes produced similar lift curves for all but two 

of the seven ice shapes tested. The two cases where considerable differences in lift behavior 

were observed were the CS22N and CS22S cases. With the exception of the 5-min rime ice 

shape, the addition of roughness resulted in lower lift coefficients at stall. The change in CL (i.e., 

rough LEWICE CL,stall - smooth LEWICE CL,stall) was as follows: -0.1, -0.03, +0.04, -0.03, -0.08, 

-0.2, and -0.08 for the CS10, IS10, SC5, CS2, CS22N, CS22S, and IPSF22 cases. With the 

exception of the CS22N and CS22S cases, the rough and smooth LEWICE ice shapes resulted in 

similar CD and CH. 

4-2
The trends in aerodynamic performance losses observed with the smooth and rough LEWICE ice 

shapes tested were consistent with results from other experimental studies involving LEWICE 

ice shapes. 

4.7 ICING RESEARCH TUNNEL VERSUS ROUGH LEWICE ICE SHAPES. 

With the exception of icing condition 5 (22.5-min complete scallop case), the rough LEWICE ice 

shapes caused similar loss in lift compared with the IRT ice shape castings. The difference in 

iced wing CL,stall defined by (ΔCL,stall = IRT casting CLstall - rough LEWICE CLstall) was -0.07, 

-0.01, -0.08, +0.05, -0.35, -0.28, and -0.08 for CS10, IS10, SC5, CS2, CS22N, CS22S, and 

IPSF22 respectively. 

In most cases, the wing drag increase caused by the LEWICE and IRT ice shapes was in good 

correlation. The difference in iced wing CD,min (i.e., Rough LEWICE CD,min - IRT ice shape 

CD,min) was 0.004, -0.008, -0.001, +0.002, -0.036, -0.013, and -0.013 for CS10, IS10, SC5, CS2, 

CS22N, CS22S, and IPSF22 respectively. 

With the exception of icing condition 5, the trends in aileron CH obtained with the rough 

LEWICE and IRT ice shapes were in good agreement. For icing conditions 2 and 3, the IRT and 

LEWICE CH magnitudes were in good correlation throughout the α-range. For icing conditions 

1, 4, and 6, the LEWICE ice shapes resulted in larger CH over regions A and B of the CH curve 

compared with the IRT ice shapes. 

In summary, the IRT and the rough LEWICE ice shapes produced similar aerodynamic effects 

for five out of the six icing conditions tested. This is an important finding considering the 

observed differences between the actual and the simulated ice shapes. It must be stressed, 

however, that considerable more experimental work is needed with a range of ice accretions to 

determine the generality of this finding. 

4.8 AILERON PERFORMANCE. 

For the clean wing and, for practically, all iced wing cases presented, the aileron remained 

effective in increasing and decreasing lift with aileron deflection. The ice shapes caused 

considerable changes to the CH of the clean wing. However, in all cases, the iced wing CH

remained within the maximum and minimum limits of the clean wing CH. 

4.9 REYNOLDS NUMBER EFFECTS. 

In general, for the low Reynolds number (Re) range of 0.5 to 1.8 million used in this study, the 

effects of Re on iced wing lift performance were small. The only notable Re effect was in the 

behavior of CL,stall for the 5-min rime ice shapes. For these ice shapes, lift near stall decreased as 

Re was increased from 1.0 to 1.8 million. 

4-3/4-4
5. REFERENCES. 

1. Lynch, F.T. and Khodadoust, A., “Effects of Ice Accretions on Aircraft Aerodynamics,” 

Progress in Aerospace Sciences, Vol. 37, Issue 8, November 2001, pp. 669-767. 

2. Vargas, M. and Reshotko, E., “Physical Mechanisms of Glaze Ice Scallop Formations on 

Swept Wings,” NASA TM-1998-206616, AIAA Paper 98-0491, January 1998. 

3. Vargas, M., “Ice Accretion on Swept Wings at Glaze Ice Conditions,” Ph.D. Thesis, Case 

Western Reserve University, Cleveland, Ohio, May 1998. 

4. Vargas, M. and Reshotko, E., “Parametric Experimental Study of the Formation of Glaze 

Ice Shapes on Swept Wings,” NASA TM-1999-208900, AIAA Paper 99-0094, January 

1999. 

5. Wright, W.B., “User Manual for the NASA Glenn Ice Accretion Code LEWICE (Version 

2.0),” NASA CR-1999-209409, September 1999. 

6. Potapczuk, M., Papadakis, M. and Vargas, M., “LEWICE Modeling of Swept Wing Ice 

Accretions,” AIAA-2003-0730, January 2003. 

7. “Multi-Hole Probe Information Packet,” Edition 1.5, Aeroprobe Corporation, October 

2001. 

8. Vargas, M., Papadakis, M., Potapczuk, M., Addy, H., Sheldon, D., and Giriunas, J., “Ice 

Accretions on a Swept GLC-305 Airfoil,” NASA TM-2002-211557, SAE Paper 2002-01-

1519, General Aviation Technology Conference and Exhibition, Wichita, KS, April 

2002. 

9. Personal communication between Dr. James T. Riley, Dr. Dave N. Anderson, and Mr. 

Harold E. Addy. Mr. Addy provided the reference icing condition. Dr. Anderson did the 

scaling from the reference condition. 

10. Addy, H., “Ice Accretions and Icing Effects for Modern Airfoils,” DOT/FAA/AR-99/89, 

NASA TP-2000-210031, April 2000. 

11. Codes of Federal Regulations, Aeronautics and Space, Part 1 to 59. 

12. Johnson, B.L., “Three-Dimensional/Reflection Place Force Data Acquisition and 

Boundary Corrections for the Walter H. Beech Memorial 7-x-10-foot Low Speed Wing 

Tunnel,” AR96-2, NIAR, Wichita State University, April 1996. 

13. Shapiro, A.H., “The Dynamics and Thermodynamics of Compressible Fluid Flow – 

Volume I,” John Wiley and Sons, 1953. 

14. Anderson, J.D., “Fundamentals of Aerodynamics,” McGraw-Hill Book Company, 1984. 

5-1
15. Drela, M., “XFOIL 6.6 User Primer,” MIT Aero and Astro Engineering, March 1996. 

16. McCormick, B.W., “Aerodynamics Aeronautics and Flight Mechanics,” 2nd edition, John 

Wiley and Sons, 1995. 

17. Addy, H.E., Potapczuk, M. and Sheldon, D.W., “Modern Airfoil Ice Accretions,” AIAA 

Paper 92-0647, January 1992. 

18. Hoerner, I.S.F. and Borst, H.V., “Fluid Dynamic – Lift,” Hoerner Fluid Dynamics, P. O. 

Box 342, Brick Town, NJ, 1975. 

19. Torenbeek, E., “Synthesis of Subsonic Airplane Design,” Delft University Press, 1982. 

20. Vargas, M. and Reshotko, E., “LWC and Temperature Effects of Ice Accretion 

Formation on Swept Wings at Glaze Ice Conditions,” AIAA Paper 2000-0483, January 

2000. 

21. Vargas, M., Giriunas, J.A. and Ratvasky, T.P., “Ice Accretion Formations on a NACA 

0012 Swept Wing Tip in Natural Icing Conditions,” AIAA Paper 2002-0244, January 

2002. 

22. Papadakis, M. Yeong, H.W., Chandrasekharan, V.R., Hinson, M. and Ratvasky, T.P., 

“Effects of Roughness on the Aerodynamic Performance of a Business Jet Tail,” AIAA 

Paper 2002-0242, January 2002. 

23. Dorsch, R.G. and Brun, R.J., “A Method for Determining Cloud-Droplet Impingement on 

Swept Wings,” NACA TN 2931, April 1953. 

5-2
Run Number Description 

AOA 

(deg.) 

V 

(mph) 

Ttotal

(°F) 

LWC 

(g/m3) 

MVD 

(μm) 

τ

(min) 

071701.01_LE_1_T Complete Scallop Condition 4 250.00 25.0 0.68 20.0 10.0 071701.02_LE_1_T Complete Scallop Condition 4 250.00 25.0 0.68 20.0 10.0 071701.03_LE_2_T Incomplete Scallop Condition 4 150.00 25.0 0.65 20.0 10.0 071701.04_LE_3_T Run based on Scaled Condition 2 for run in 2D table 6 201.30 11.7 0.51 14.5 5.0 071701.05_LE_2_T SLD Condition A 4 172.60 25.0 0.60 70.0 10.0 071701.06_LE_2_T Incomplete Scallop Condition 4 150.00 25.0 0.65 20.0 10.0 071801.01_LE_1_T SLD Condition C 4 172.62 25.0 1.00 175.0 22.5 071801.02_LE_1_T 22.5 minutes failed ice protection App C 4 150.00 27.0 0.46 20.0 22.5 071801.03_LE_2_T Complete Scallop Condition Long ice accretion time 4 250.00 25.0 0.68 20.0 22.5 071801.04_LE_3_T SLD Condition D 4 172.62 30.0 1.00 175.0 15.0 071801.05_LE_3_T Complete Scallop Condition Short ice accretion time 4 250.00 25.0 0.68 20.0 2.0 

071901.01_LE_1_M 

071901.01_LE_1_B SLD Condition B 4 172.60 25.0 1.00 175.0 22.5 

071901.02_LE_1_M 

071901.02_LE_1_B 

22.5 minutes failed ice protection 

App C 4 150.00 27.0 0.46 20.0 22.5 

071901.03_LE_2_M 

071901.03_LE_2_B 

Complete Scallop Condition 

Long ice accretion time 4 250.00 25.0 0.68 20.0 22.5 

071901.04_LE_2_M 

071901.04_LE_2_B Incomplete Scallop Condition 4 150.00 25.0 0.65 20.0 10.0 

071901.05_LE_3_M 

071901.05_LE_3_B Complete Scallop Condition 4 250.00 25.0 0.68 20.0 10.0 

072001.01_LE_1_M 

072001.01_LE_1_B 

Run based on Scaled Condition 2 

for run in 2D table 6 201.30 11.7 0.51 14.5 5.0 

072001.02_LE_1_M 

072001.02_LE_1_B 

Run based on Scaled Condition 2 

for run in 2D table 6 201.30 11.7 0.51 14.5 5.0 

072001.03_LE_2_M 

072001.03_LE_2_B 

Complete Scallop Condition 

Short ice accretion time 4 250.00 25.0 0.68 20.0 2.0 

072001.04_LE_3_M 

072001.04_LE_3_B 

Complete Scallop Condition 

Long ice accretion time 4 250.00 25.0 0.68 20.0 22.5 

072001.05_LE_3_M 

072001.05_LE_3_B 

Substitute condition to avoid 

losing the ice accretion 4 225.00 25.0 0.68 20.0 22.5 

APPENDIX A—RUN LOG FOR ICING TEST AT NASA GLENN ICING

RESEARCH TUNNEL 

A-1/A-2
APPENDIX B—RUN LOG FOR AERODYNAMIC INVESTIGATION AT 

WICHITA STATE UNIVERSITY 

Date Run No. δA (deg) Configuration Name Comments 

25-Apr-02 1 0 Clean 

2 -15 Clean 

3 -10 Clean 

4 -5 Clean 

5 -2.5 Clean 

6 0 Clean 

7 2.5 Clean 

8 5 Clean 

9 10 Clean 

10 15 Clean 

11 20 Clean 

26-Apr-02 13 20 IRT-CS10 (irt-ice1) Repeat R012 

14 15 IRT-CS10 (irt-ice1) 

15 10 IRT-CS10 (irt-ice1) 

16 5 IRT-CS10 (irt-ice1) 

17 2.5 IRT-CS10 (irt-ice1) 

18 0 IRT-CS10 (irt-ice1) 

19 -2.5 IRT-CS10 (irt-ice1) 

20 -5 IRT-CS10 (irt-ice1) 

21 -10 IRT-CS10 (irt-ice1) 

22 -15 IRT-CS10 (irt-ice1) 

23 0 IRT-CS10 (irt-ice1) 

24 0 Clean 

29-Apr-02 26 0 Clean Gap opened 

27 0 Clean Gap closed 

28 -2.5 IRT-IS10 (irt-ice2) 

29 -5 IRT-IS10 (irt-ice2) 

30 -10 IRT-IS10 (irt-ice2) 

31 -15 IRT-IS10 (irt-ice2) 

32 0 IRT-IS10 (irt-ice2) 

33 2.5 IRT-IS10 (irt-ice2) 

34 5 IRT-IS10 (irt-ice2) Ice shape broke 

30-Apr-02 36 0 Clean Baseline repeat 

37 -15 IRT-CS22 (irt-ice5) 

38 -10 IRT-CS22 (irt-ice5) 

39 -5 IRT-CS22 (irt-ice5) 

40 -2.5 IRT-CS22 (irt-ice5) 

41 0 IRT-CS22 (irt-ice5) 

42 2.5 IRT-CS22 (irt-ice5) 

43 5 IRT-CS22 (irt-ice5) 

B-1
Date Run No. δA (deg) Configuration Name Comments 

1-May-02 44 10 IRT-CS22 (irt-ice5) 

45 15 IRT-CS22 (irt-ice5) 

46 20 IRT-CS22 (irt-ice5) 

47 0 IRT-CS22 (irt-ice5) 

49 0 Clean Baseline repeat 

50 0 IRT-SC5 (irt-ice3) 

51 -15 IRT-SC5 (irt-ice3) 

52 -10 IRT-SC5 (irt-ice3) 

53 -5 IRT-SC5 (irt-ice3) 

54 -2.5 IRT-SC5 (irt-ice3) 

55 0 IRT-SC5 (irt-ice3) 

2-May-02 56 2.5 IRT-SC5 (irt-ice3) Loose tape 

57 2.5 IRT-SC5 (irt-ice3) Loose tape 

58 2.5 IRT-SC5 (irt-ice3) Repeat R056 and R057 

59 5 IRT-SC5 (irt-ice3) 

60 10 IRT-SC5 (irt-ice3) 

61 15 IRT-SC5 (irt-ice3) 

62 20 IRT-SC5 (irt-ice3) 

64 0 Clean Baseline repeat 

65 0 IRT-CS2 (irt-ice4) 

66 -15 IRT-CS2 (irt-ice4) 

67 -10 IRT-CS2 (irt-ice4) 

68 -5 IRT-CS2 (irt-ice4) 

69 -2.5 IRT-CS2 (irt-ice4) 

3-May-02 70 0 IRT-CS2 (irt-ice4) 

71 2.5 IRT-CS2 (irt-ice4) 

72 5 IRT-CS2 (irt-ice4) 

73 10 IRT-CS2 (irt-ice4) 

74 15 IRT-CS2 (irt-ice4) 

75 20 IRT-CS2 (irt-ice4) 

82 0 Clean Baseline repeat 

6-May-02 84 0 Clean Baseline repeat 

85 0 IRT-IPSF22 (irt-ice6) 

86 -15 IRT-IPSF22 (irt-ice6) 

87 -10 IRT-IPSF22 (irt-ice6) 

88 -5 IRT-IPSF22 (irt-ice6) 

89 -5 IRT-IPSF22 (irt-ice6) Repeat R088 

90 -2.5 IRT-IPSF22 (irt-ice6) 

91 0 IRT-IPSF22 (irt-ice6) 

92 2.5 IRT-IPSF22 (irt-ice6) 

93 5 IRT-IPSF22 (irt-ice6) Max out speed; Ttunnel = 152°F 

94 10 IRT-IPSF22 (irt-ice6) 

B-2
Date Run No. δA (deg) Configuration Name Comments 

7-May-02 95 15 IRT-IPSF22 (irt-ice6) 

96 20 IRT-IPSF22 (irt-ice6) Loose tape. Terminate run 

97 20 IRT-IPSF22 (irt-ice6) Repeat R096 

99 0 Clean Baseline repeat (ReMAC = 1.8x106) 

100 0 Clean ReMAC = 1.5x106

101 0 Clean ReMAC = 1.0x106

102 0 Clean ReMAC = 0.5x106

103 0 IRT-IS10 (irt-ice2) Repeat R032 

104 2.5 IRT-IS10 (irt-ice2) Repeat R033 

105 5 IRT-IS10 (irt-ice2) Repeat R034 

106 10 IRT-IS10 (irt-ice2) 

107 15 IRT-IS10 (irt-ice2) 

8-May-02 108 20 IRT-IS10 (irt-ice2) 

109 -15 IRT-IS10 (irt-ice2) 

110 0 IRT-IS10 (irt-ice2) 

111 0 IRT-CS22 (irt-ice5) ReMAC = 0.5x106

112 0 IRT-CS22 (irt-ice5) ReMAC = 1.5x106

113 0 IRT-CS22 (irt-ice5) ReMAC = 1.0x106

114 0 IRT-CS22 (irt-ice5) ReMAC = 1.8x106

115 0 IRT-SC5 (irt-ice3) ReMAC = 1.8x106 (tape loose) 

116 0 IRT-SC5 (irt-ice3) ReMAC = 1.8x106; Repeat R115 

117 0 IRT-SC5 (irt-ice3) ReMAC = 1.5x106

118 0 IRT-SC5 (irt-ice3) ReMAC = 1.0x106

119 0 IRT-SC5 (irt-ice3) ReMAC = 0.5x106

121 0 Clean Baseline repeat 

9-May-02 135 0 Clean Baseline repeat 

136 0 Clean Baseline (using Model Constants Table 2). 

10-May-02 149 0 Clean Baseline repeat 

13-May-02 151 0 Clean Baseline repeat 

14-May-02 167 0 LS-IS10 (ls-ice2) 

168 -15 LS-IS10 (ls-ice2) 

169 -10 LS-IS10 (ls-ice2) 

170 -5 LS-IS10 (ls-ice2) 

171 -2.5 LS-IS10 (ls-ice2) 

172 0 LS-IS10 (ls-ice2) 

173 2.5 LS-IS10 (ls-ice2) 

174 5 LS-IS10 (ls-ice2) 

15-May-02 175 10 LS-IS10 (ls-ice2) 

176 15 LS-IS10 (ls-ice2) 

177 20 LS-IS10 (ls-ice2) 

178 0 LS-IS10 (ls-ice2) 

180 0 Clean Baseline repeat 

181 0 LS-SC5 (ls-ice3) 

B-3
Date Run No. δA (deg) Configuration Name Comments 

15-May-02 182 -15 LS-SC5 (ls-ice3) 

183 -10 LS-SC5 (ls-ice3) 

184 -5 LS-SC5 (ls-ice3) 

185 -2.5 LS-SC5 (ls-ice3) 

186 0 LS-SC5 (ls-ice3) 

187 0 LS-SC5 (ls-ice3) 

16-May-02 188 2.5 LS-SC5 (ls-ice3) 

189 5 LS-SC5 (ls-ice3) 

190 10 LS-SC5 (ls-ice3) 

191 15 LS-SC5 (ls-ice3) 

192 20 LS-SC5 (ls-ice3) 

193 0 LS-SC5 (ls-ice3) 

195 0 Clean Baseline repeat 

196 0 LR-IS10 (lr-ice2) 

197 -15 LR-IS10 (lr-ice2) 

198 0 LR-IS10 (lr-ice2) Repeat of R196 (start at α = 0°), losing grit 

199 -10 LR-IS10 (lr-ice2) 

200 -5 LR-IS10 (lr-ice2) 

17-May-02 201 0 LS-CS2 (ls-ice4) 

202 -15 LS-CS2 (ls-ice4) 

203 -10 LS-CS2 (ls-ice4) 

204 -5 LS-CS2 (ls-ice4) 

205 -2.5 LS-CS2 (ls-ice4) 

206 0 LS-CS2 (ls-ice4) 

207 2.5 LS-CS2 (ls-ice4) 

208 5 LS-CS2 (ls-ice4) 

209 10 LS-CS2 (ls-ice4) 

210 15 LS-CS2 (ls-ice4) 

211 20 LS-CS2 (ls-ice4) 

213 0 LR-SC5 (lr-ice3) 

214 -15 LR-SC5 (lr-ice3) 

215 -10 LR-SC5 (lr-ice3) 

20-May-02 216 0 LR-SC5 (lr-ice3) 

217 -5 LR-SC5 (lr-ice3) 

218 -2.5 LR-SC5 (lr-ice3) 

219 2.5 LR-SC5 (lr-ice3) 

220 5 LR-SC5 (lr-ice3) 

221 10 LR-SC5 (lr-ice3) 

222 15 LR-SC5 (lr-ice3) 

223 20 LR-SC5 (lr-ice3) 

225 0 Clean Baseline repeat 

226 0 LR-CS2 (lr-ice4) Use static tare R1114 

B-4
Date Run No. δA (deg) Configuration Name Comments 

20-May-02 227 -15 LR-CS2 (lr-ice4) 

228 -10 LR-CS2 (lr-ice4) 

229 -5 LR-CS2 (lr-ice4) 

230 -2.5 LR-CS2 (lr-ice4) 

231 2.5 LR-CS2 (lr-ice4) 

21-May-02 232 0 LR-CS2 (lr-ice4) 

233 2.5 LR-CS2 (lr-ice4) 

234 5 LR-CS2 (lr-ice4) 

235 10 LR-CS2 (lr-ice4) 

236 15 LR-CS2 (lr-ice4) 

237 20 LR-CS2 (lr-ice4) 

239 0 Clean Baseline repeat 

22-May-02 240 0 LR-IS10 (lr-ice2) Use static tare R1113 

241 -5 LR-IS10 (lr-ice2) Repeat of R199 

242 -2.5 LR-IS10 (lr-ice2) 

243 2.5 LR-IS10 (lr-ice2) 

244 5 LR-IS10 (lr-ice2) 

245 10 LR-IS10 (lr-ice2) 

246 15 LR-IS10 (lr-ice2) 

247 20 LR-IS10 (lr-ice2) 

24-May-02 264 0 Clean Baseline repeat 

28-May-02 276 0 Clean Baseline repeat 

29-May-02 291 0 Clean Baseline repeat 

30-May-02 299 0 Clean Baseline repeat 

31-May-02 314 0 Clean Baseline repeat 

315 -15 Clean (fixed transition) 2 layers of aluminum strips at 2%c 

316 -5 Clean (fixed transition) 2 layers of aluminum strips at 2%c 

317 0 Clean (fixed transition) 2 layers of aluminum strips at 2%c 

318 5 Clean (fixed transition) 2 layers of aluminum strips at 2%c 

319 15 Clean (fixed transition) 2 layers of aluminum strips at 2%c 

3-Jun-02 325 0 Clean Baseline repeat 

328 0 LS-CS10 (ls-ice1) 

329 -15 LS-CS10 (ls-ice1) 

330 -10 LS-CS10 (ls-ice1) 

331 -5 LS-CS10 (ls-ice1) 

332 -2.5 LS-CS10 (ls-ice1) 

333 2.5 LS-CS10 (ls-ice1) 

4-Jun-02 334 0 LS-CS10 (ls-ice1) 

335 5 LS-CS10 (ls-ice1) 

336 10 LS-CS10 (ls-ice1) 

337 15 LS-CS10 (ls-ice1) 

338 20 LS-CS10 (ls-ice1) 

B-5
Date Run No. δA (deg) Configuration Name Comments 

4 June-02 339 0 LS-IPSF22 (ls-ice6) 

340 -15 LS-IPSF22 (ls-ice6) 

341 -10 LS-IPSF22 (ls-ice6) 

342 -5 LS-IPSF22 (ls-ice6) 

343 -2.5 LS-IPSF22 (ls-ice6) 

344 2.5 LS-IPSF22 (ls-ice6) 

345 5 LS-IPSF22 (ls-ice6) 

346 10 LS-IPSF22 (ls-ice6) 

347 15 LS-IPSF22 (ls-ice6) 

5-Jun-02 348 0 LS-IPSF22 (ls-ice6) 

349 15 LS-IPSF22 (ls-ice6) Repeat R347 

350 20 LS-IPSF22 (ls-ice6) 

351 10 LS-IPSF22 (ls-ice6) Repeat R346 

352 0 LR-CS10 (lr-ice1) Use Static Tare R1116 

353 -15 LR-CS10 (lr-ice1) 

354 -10 LR-CS10 (lr-ice1) 

355 -5 LR-CS10 (lr-ice1) 

356 -2.5 LR-CS10 (lr-ice1) 

357 2.5 LR-CS10 (lr-ice1) 

358 5 LR-CS10 (lr-ice1) 

359 10 LR-CS10 (lr-ice1) 

360 15 LR-CS10 (lr-ice1) 

361 20 LR-CS10 (lr-ice1) Aileron binding. Terminate run 

362 20 LR-CS10 (lr-ice1) Repeat R361 

363 0 LR-CS10 (lr-ice1) 

6-Jun-02 364 0 LR-IPSF22 (lr-ice6) 

365 20 LR-IPSF22 (lr-ice6) 

366 15 LR-IPSF22 (lr-ice6) 

367 10 LR-IPSF22 (lr-ice6) 

368 5 LR-IPSF22 (lr-ice6) 

369 2.5 LR-IPSF22 (lr-ice6) 

370 -2.5 LR-IPSF22 (lr-ice6) 

371 -5 LR-IPSF22 (lr-ice6) 

372 -10 LR-IPSF22 (lr-ice6) 

373 -15 LR-IPSF22 (lr-ice6) 

374 0 LR-IPSF22 (lr-ice6) 

376 0 Clean Baseline repeat 

10-Jun-02 390 0 Clean Baseline repeat 

11-Jun-02 396 0 Clean Baseline repeat (automation wrong) 

397 0 Clean Baseline repeat 

399 0 IRT-CS22 (irt-ice5) 

400 0 IRT-IS10 (irt-ice2) 

B-6
Date Run No. δA (deg) Configuration Name Comments 

12-Jun-02 402 0 Clean Baseline repeat 

403 0 LS-CS22N (ls-ice5n) 

404 20 LS-CS22N (ls-ice5n) 

405 15 LS-CS22N (ls-ice5n) 

406 10 LS-CS22N (ls-ice5n) 

407 5 LS-CS22N (ls-ice5n) 

Horn on pressure side split, due to heat. 

Was glued back and reinforced with 

screws. 

408 5 LS-CS22N (ls-ice5n) Repeat of R407 

409 2.5 LS-CS22N (ls-ice5n) 

410 -2.5 LS-CS22N (ls-ice5n) 

411 -5 LS-CS22N (ls-ice5n) 

412 -10 LS-CS22N (ls-ice5n) 

13-Jun-02 413 -15 LS-CS22N (ls-ice5n) 

414 0 LS-CS22N (ls-ice5n) 

417 0 LS-CS22N (ls-ice5n) ReMAC = 0.5x106

418 0 LS-CS22N (ls-ice5n) ReMAC = 1.0x106

419 0 LS-CS22N (ls-ice5n) ReMAC = 1.5x106

420 0 LR-IPSF22 (lr-ice6) Repeat R364 and R374 

421 0 LR-IPSF22 (lr-ice6) ReMAC = 1.5x106

422 0 LR-IPSF22 (lr-ice6) ReMAC = 1.0x106

423 0 LR-IPSF22 (lr-ice6) ReMAC = 0.5x106

424 0 LR-IS10 (lr-ice2) Repeat R196 and R240 

425 0 LR-SC5 (lr-ice3) Repeat R213 and R216 

14-Jun-02 426 0 LR-SC5 (lr-ice3) Re-installed ice shape. Repeat R425 

427 0 LR-SC5 (lr-ice3) ReMAC = 1.5x106

428 0 LR-SC5 (lr-ice3) ReMAC = 1.0x106

429 0 LR-SC5 (lr-ice3) ReMAC = 0.5x106

430 0 LS-CS22S (ls-ice5s) 

431 0 LS-CS22S (ls-ice5s) 

432 20 LS-CS22S (ls-ice5s) 

433 15 LS-CS22S (ls-ice5s) 

434 10 LS-CS22S (ls-ice5s) 

435 5 LS-CS22S (ls-ice5s) 

436 2.5 LS-CS22S (ls-ice5s) 

437 -2.5 LS-CS22S (ls-ice5s) 

438 -5 LS-CS22S (ls-ice5s) 

439 -10 LS-CS22S (ls-ice5s) Aileron binding. Terminate run 

17-Jun-02 440 -10 LS-CS22S (ls-ice5s) Repeat R439 

441 -15 LS-CS22S (ls-ice5s) 

442 0 LS-CS22S (ls-ice5s) 

444 0 Clean Baseline repeat 

445 0 LR-CS22N (lr-ice5n) 

B-7
Date Run No. δA (deg) Configuration Name Comments 

17-Jun-02 446 20 LR-CS22N (lr-ice5n) 

447 15 LR-CS22N (lr-ice5n) 

448 10 LR-CS22N (lr-ice5n) 

449 5 LR-CS22N (lr-ice5n) 

450 2.5 LR-CS22N (lr-ice5n) 

451 -2.5 LR-CS22N (lr-ice5n) Loose tape. Terminate run 

18-Jun-02 455 -2.5 LR-CS22N (lr-ice5n) Repeat R451 

456 -5 LR-CS22N (lr-ice5n) 

457 -10 LR-CS22N (lr-ice5n) 

458 -15 LR-CS22N (lr-ice5n) 

459 0 LR-CS22N (lr-ice5n) 

460 0 LR-CS22S (lr-ice5s) 

461 20 LR-CS22S (lr-ice5s) 

462 15 LR-CS22S (lr-ice5s) 

463 10 LR-CS22S (lr-ice5s) 

464 5 LR-CS22S (lr-ice5s) Loose tape. Terminate run 

465 0 LR-CS22S (lr-ice5s) Repeat R460, loose tape 

466 0 LR-CS22S (lr-ice5s) Repeat R460, loose tape 

467 0 LR-CS22S (lr-ice5s) Repeat R460, loose tape 

19-Jun-02 468 0 LR-CS22S (lr-ice5s) Repeat R460 

469 5 LR-CS22S (lr-ice5s) Repeat R464 

470 2.5 LR-CS22S (lr-ice5s) 

471 -2.5 LR-CS22S (lr-ice5s) Sandpaper started to peel off at about 17°, 

but data looks good. 

472 -5 LR-CS22S (lr-ice5s) 

473 -10 LR-CS22S (lr-ice5s) 

474 -15 LR-CS22S (lr-ice5s) Lost sandpaper 

475 -15 LR-CS22S (lr-ice5s) Repeat R474 

476 0 LR-CS22S (lr-ice5s) ReMAC = 1.5x106

477 0 LR-CS22S (lr-ice5s) ReMAC = 1.0x106

478 0 LR-CS22S (lr-ice5s) ReMAC = 0.5x106

480 15 Clean Repeat R010 

481 0 LR-CS22N (lr-ice5n) Repeat R445 

482 0 LR-CS22N (lr-ice5n) ReMAC = 0.5x106

20-Jun-02 483 0 LR-CS22N (lr-ice5n) ReMAC = 1.0x106

484 0 LR-CS22N (lr-ice5n) ReMAC = 1.5x106

485 0 LR-CS10 (lr-ice1) Repeat R352 

486 0 LR-CS10 (lr-ice1) ReMAC = 1.5x106

487 0 LR-CS10 (lr-ice1) ReMAC = 1.0x106

488 0 LR-CS10 (lr-ice1) ReMAC = 0.5x106

21-Jun-02 501 0 Clean Baseline repeat 

25-Jun-02 521 0 Clean Flow-viz, ReMAC = 1.5x106

522 0 IRT-CS10 (irt-ice1) Flow-viz, ReMAC = 1.5x106

B-8
Date Run No. δA (deg) Configuration Name Comments 

25-Jun-02 523 0 IRT-IPSF22 (irt-ice6) Flow-viz, ReMAC = 1.5x106

524 0 IRT-IS10 (irt-ice2) Flow-viz, ReMAC = 1.5x106

525 0 IRT-SC5 (irt-ice3) Flow-viz, ReMAC = 1.5x106

526 0 IRT-CS2 (irt-ice4) Flow-viz, ReMAC = 1.5x106

26-Jun-02 527 0 LR-IS10 (lr-ice2) Flow-viz, ReMAC = 1.5x106

528 0 LR-SC5 (lr-ice3) Flow-viz, ReMAC = 1.5x106

529 0 LR-CS2 (lr-ice4) Flow-viz, ReMAC = 1.5x106

530 0 LR-CS10 (lr-ice1) Flow-viz, ReMAC = 1.5x106

531 0 LR-IPSF22 (lr-ice6) Flow-viz, ReMAC = 1.5x106

532 0 LR-CS22S (lr-ice5s) Flow-viz, ReMAC = 1.5x106

533 0 LR-CS22N (lr-ice5n) Flow-viz, ReMAC = 1.5x106

Repeat R520, ReMAC = 1.0x106 534 0 IRT-CS22 (irt-ice5) 

Repeat R521, ReMAC = 1.5x106 536 0 Clean 

Flow-viz, ReMAC = 1.5x106 27-Jun-02 537 20 Clean 

Flow-viz, ReMAC = 1.5x106 538 -15 Clean 

27-Jun-02 539 0 Clean Take data with yarn tufts on model, 

ReMAC = 1.5x106

541 0 Clean Baseline repeat 

548 0 LR-IPSF22 (lr-ice6) Added 24-grit to tip of horns of ice shape 

549 15 LR-IPSF22 (lr-ice6) Added 24-grit to tip of horns of ice shape 

550 -15 LR-IPSF22 (lr-ice6) Added 24-grit to tip of horns of ice shape 

1-Jul-02 564 0 IRT-IPSF22 (irt-ice6) ReMAC = 1.8x106

565 0 IRT-IPSF22 (irt-ice6) ReMAC = 1.5x106

566 0 IRT-IPSF22 (irt-ice6) ReMAC = 1.0x106

567 0 IRT-IPSF22 (irt-ice6) ReMAC = 0.5x106

568 0 LS-IPSF22 (ls-ice6) ReMAC = 1.8x106

1-Jul-02 569 0 LS-IPSF22 (ls-ice6) ReMAC = 1.5x106

570 0 LS-IPSF22 (ls-ice6) ReMAC = 1.0x106

571 0 LS-IPSF22 (ls-ice6) ReMAC = 0.5x106

573 0 Clean Baseline repeat 

2-Jul-02 582 0 LS-CS22S (ls-ice5s) Flow-viz, ReMAC = 1.5x106

, one of the VCR 

was not recording 

583 0 LS-CS22S (ls-ice5s) Flow-viz, ReMAC = 1.5x106, Repeat R582 

584 0 LS-IPSF22 (ls-ice6) Flow-viz, ReMAC = 1.5x106

3-Jul-02 587 0 IRT-CS22 (irt-ice5) Flow-viz, ReMAC = 1.5x106

5-Jul-02 595 0 Clean Baseline repeat 

Checking flow angularity probe installation, 

q = 15 psf 596 0 Clean 

597 0 Clean 7-hole flow angularity probes 

598 20 Clean 7-hole flow angularity probes 

599 -15 Clean 7-hole flow angularity probes 

603 0 Clean Gap opened 

604 0 IRT-CS22 (irt-ice5) Gap opened 

B-9/B-10
APPENDIX C—PRESSURE DATA FROM ICING TESTS

Surface pressures at 15%, 50%, and 85% semispan locations were obtained prior to the ice 

accretion tests. At velocity of approximately 150 mph, with the wing in normal position (refer to 

section 2.1.3), pressures were measured at angles of attack of -3°, -2.5°, -2°, -1.5°, -1°, -0.5°, 0°, 

0.5°, 1°, 1.5°, 2°, 4°, 6°, 8°, 10°, 11°, 12°, 13°, 14°, and 14.5°. Pressure coefficient (Cp) 

distributions of selected α are presented in figure C-1. 

Pressure measurements were performed with the icing research tunnel (IRT) electronically 

scanned pressure (ESP) system. Six 32-port (±5 psid) ESP modules were available in the IRT, 

providing a total of 192 pressure channels. One port in each module was used for a check 

pressure; thus 31 channels per module were available for test data, or a total of 186 ports. The 

ESP system applied a three-point pressure calibration to all port transducers. This on-line threepoint calibration ensured that measurement errors were not greater than ±0.1% of full-scale. The 

standard calibration interval was every 400 cycles (approximately 15 minutes). 

The experimental pressure data were used to validate Cp distributions from LEWICE analyses 

and wind tunnel tests at Wichita State University. 

C-1
(a) Cp vs. x/c, α = -3°

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

-2.5

-3.0

-3.5

Pressure coefficient

α = -3 deg

15%-semispan

50%-semispan

85%-semispan

(b) Cp vs. x/c, α = 0°

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

Pressure coefficient

α = 0 deg

15%-semispan

50%-semispan

85%-semispan

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

-2.5

-3.0

-3.5

-4.0

-4.5

Pressure coefficient

α = 10 deg

15%-semispan

50%-semispan

85%-semispan

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

-2.5

Pressure coefficient

α = 6 deg

15%-semispan

50%-semispan

85%-semispan

(c) Cp vs. x/c, α = 6° (d) Cp vs. x/c, α = 10°

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

-2.5

-3.0

-3.5

-4.0

-4.5

-5.0

Pressure coefficient

α = 12 deg

15%-semispan

50%-semispan

85%-semispan

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0

X/C

1.0

0.5

0.0

-0.5

-1.0

-1.5

-2.0

-2.5

-3.0

Pressure coefficient

α = 14 deg

15%-semispan

50%-semispan

85%-semispan

(e) Cp vs. x/c, α = 12° (f) Cp vs. x/c, α = 14°

FIGURE C-1. PRESSURE DISTRIBUTIONS OF CLEAN CONFIGURATION FROM IRT; 

WING IN NORMAL POSITION; V∞ = 150 mph 

C-2
APPENDIX D—FLOW ANGULARITY STUDIES

Two, seven-hole flow angularity probes were installed on the clean wing at two stations 

corresponding to approximately 16% and 80% semispan. Seven-hole probes can measure the 

three components of velocity, the total pressure, and the static pressure at a point in the flow. 

The flow probe provides results with high accuracy for angles as high as 75°. The data from the 

seven-hole probes were used to estimate the local angle of attack of the wing due to geometric 

twist of the wing, presence of the streamlined body, and vortices at the wing tip. Angle of attack 

sweeps were conducted for three aileron deflections (-15°, 0°, and 20°) and at Reynolds number 

of 1.8 million based on wing mean aerodynamic chord. Pressures measured from the seven-hole 

probes were recorded. Installation of the seven-hole flow probes is shown in figures 2-52 to 

2-55. 

In order for the probes to provide accurate downwash angle measurements, they were attached 

onto the wing leading edge and aligned to the free-stream flow direction. Note that this line of 

direction is not concurrent to the wing chordwise direction, and therefore, a set of brackets was 

specifically made to accommodate the installation. 

Effects of flow probes on the wing lift and drag coefficients are demonstrated in figures D-1 and 

D-2. It is observed that the flow probes improved lift performance of the wing; i.e., lift 

coefficient (CL) was increased throughout the α-range and as a result CL,stall was increased as 

well. This is because the flow probes generated vortices that re-energized the boundary layer, 

which delayed transition and flow separation; hence, improved CL,stall. From figure D-1, it is 

observed that the offsets between the CL values obtained with and without the flow probes, 

amongst all three aileron deflections of -15°, 0°, and 20°, were very similar, and that higher lift 

curves were produced with higher δA. It is of interest to draw attention to the fact that the lift 

curve generated with the flow probes at δA of -15° behaved very much like that from without the 

probes at δA of 0°. 

The same vortices that re-energized the boundary layer also inadvertently increased drag 

coefficient (CD), as shown in figure D-2. In general, the increment in drag contributed by the 

probes was only accountable when the wing was experiencing a positive α. Consider the case of 

drag coefficients generated by the wing with the flow probes installed at δA of 20°, its CD values 

were the highest amongst all three aileron deflection configurations, with and without the probes, 

yet only at the range of positive α from 2° onwards. In contrary, its CD values from α of -4° to 

-8° were the lowest amongst all configurations tested. 

Downwash (or upwash) angle as a function of geometric angle of attack is depicted in 

figures D-3 and D-4. Results from figure D-3 were collected from the probe at 16% semispan 

location, whereas those from figure D-4 were collected from the one at 80% semispan. In both 

cases, the most negative aileron deflection (δA = -15°) configuration generated the largest 

amount of downwash angles for the positive α-sweep. Under the same argument, the test 

configuration of the most positive aileron deflection (δA = 20°) generated the largest upwash 

angles (negative downwash angles) during the negative α-sweep. Note that the downwash angle 

curve in figure D-4 obtained near the tip of the semispan behaved less linearly than the one 

D-1
observed towards the wing root, as shown in figure D-3. Also, the slope of the downwash angle 

with respect to the angle of attack observed at the near-tip section is smaller than that of the nearroot section. 

0.00

0.05

0.10

0.15

0.20

0.25

0.30

0.35

FIGURE D-1. EFFECTS OF FLOW PROBES 

ON CL; Re = 1.8x106

FIGURE D-2. EFFECTS OF FLOW PROBES 

ON CD; Re = 1.8x106

FIGURE D-3. RELATIONSHIP OF 

DOWNWASH ANGLE WITH GEOMETRIC 

ANGLE OF ATTACK; 16% SEMISPAN; 

Re = 1.8x106

FIGURE D-4. RELATIONSHIP OF 

DOWNWASH ANGLE WITH GEOMETRIC 

ANGLE OF ATTACK; 80% SEMISPAN; 

Re = 1.8x106

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Lift coefficient

No flow probe, δA = 0 deg

No flow probe, δA = 20 deg

No flow probe, δA = -15 deg

With flow probe, δA = 0 deg

With flow probe, δA = 20 deg

With flow probe, δA = -15 deg

No flow probe, δA = 0 deg

No flow probe, δA = 20 deg

No flow probe, δA = -15 deg

With flow probe, δA = 0 deg

With flow probe, δA = 20 deg

With flow probe, δA = -15 deg

icient ff ag coe

Dr

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

Negative downwash angle 

is upwash 

-10 -5 0 5 10 15

Angle of attack (deg)

-4

-2

0

2

4

6

8

Downwash angle (deg)

δA = 0 deg (R597)

δA = 20 deg (R598)

δA = -15 deg (R599)

Negative downwash angle 

is upwash 

-10 -5 0 5 10 15

Angle of attack (deg)

-4

-2

0

2

4

6

8

Downwash angle (deg)

δA = 0 deg (R597)

δA = 20 deg (R598)

δA = -15 deg (R599)

D-2
APPENDIX E—SIMULATED FROST USING SANDPAPER

As per Federal Aviation Administration request, a total of eight configurations of frost 

simulation using 2-grit sandpaper (times four coverage) were identified and investigated to study 

the effects of frost on aerodynamic performance of a swept wing. The use of sandpaper in 

simulating ice shape roughness aligns with current procedures used by airframers during aircraft 

icing certification*

. 

Sandpaper tested were 40 grit and 150 grit. With mean aerodynamic chord (MAC) of 1.25 ft, 

40-grit provided normalized roughness (k/MAC) of 1x10-3, where the 150-grit sandpaper 

provided k/MAC of 2x10-4. Figure E-1 shows the four different frost simulation coverage using 

40-grit sandpaper. The first configuration (figure E-1(a)) simulated frost coverage on the whole 

wing. From figures E-1(b) and E-1(c), the second and third configurations were to simulate frost 

coverage on the aft 87.5% and 35% of the wing, respectively. In addition, the fourth coverage 

(figure E-1(d)) was to simulate failed deicing fluid condition. All sandpaper coverage were 

applied on the wing upper surface only. 

Figure E-2 demonstrates the effects of simulated frost on CL, CD, CM, and CH of the wing using 

40-grit sandpaper, whereas the simulated frost effect using 150-grit roughness was shown in 

figure E-3. 

 

*

 Papadakis, M., Yeong, H.W., Chadrasekharan, V.R., Hinston, M., and Ratvasky, T.P., “Effects of Roughness on 

the Aerodynamic Performance of a Business Jet Tail,” AIAA Paper 2002-0242, January 2002. 

E-1
FR1-40 (configuration 5) FR2-40 (configuration 6) 

FR3-40 (configuration 7) FR4-40 (configuration 8) 

FIGURE E-1. GLC-305 MODEL WITH 40-GRIT (K/MAC = 1X10-3) SANDPAPER, 

SIMULATING FROST FORMATION 

E-2
(a) Lift coefficient (b) Drag coefficient 

(c) Pitching-moment coefficient (d) Hinge-moment coefficient 

FIGURE E-2. EFFECT OF FROST ON AERODYNAMIC PERFORMANCE; 40-GRIT 

SANDPAPER; Re = 1.8x106

; δA = 0°

-10 -5 0 5 10 15 20 2

Angle of attack (deg)

5

0.00

0.05

0.10

0.15

0.20

0.25

0.30

0.35

Drag Coefficient

Clean (R001)

FR1-40 (R252)

FR2-40 (R253)

FR3-40 (R249)

FR4-40 (R254)

-10 -5 0 5 10 15 20 2

Angle of attack (deg)

5

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

1.2

Lift Coefficient

Clean (R001)

FR1-40 (R252)

FR2-40 (R253)

FR3-40 (R249)

FR4-40 (R254)

-0.10

-0.08

-0.06

-0.04

-0.02

0.00

0.02

Pitching Moment Coe

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

fficient

Clean (R001)

FR1-40 (R252)

FR2-40 (R253)

FR3-40 (R249)

FR4-40 (R254)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.15

-0.10

-0.05

0.00

0.05

0.10

Hinge Moment Coefficient

Clean (R001)

FR1-40 (R252)

FR2-40 (R253)

FR3-40 (R249)

FR4-40 (R254)

E-3
-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

1.2

(a) Lift coefficient (b) Drag coefficient 

(c) Pitching-moment coefficient (d) Hinge-moment coefficient 

FIGURE E-3. EFFECT OF FROST ON AERODYNAMIC PERFORMANCE; 150-GRIT 

SANDPAPER; Re = 1.8x106

; δA = 0°

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

Lift Coefficient

Clean (R001)

FR1-150 (R260)

FR2-150 (R257)

FR3-150 (R262)

FR4-150 (R261)

-10 -5 0 5 10 15 20 2

Angle of attack (deg)

5

0.00

0.05

0.10

0.15

0.20

0.25

0.30

0.35

Drag Coefficient

Clean (R001)

FR1-150 (R260)

FR2-150 (R257)

FR3-150 (R262)

FR4-150 (R261)

-10 -5 0 5 10 15 20 2

Angle of attack (deg)

5

-0.10

-0.08

-0.06

-0.04

-0.02

0.00

0.02

Pitching Moment Coefficient

Clean (R001)

FR1-150 (R260)

FR2-150 (R257)

FR3-150 (R262)

FR4-150 (R261)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.15

-0.10

-0.05

0.00

0.05

0.10

Hinge Moment Coefficient

Clean (R001)

FR1-150 (R260)

FR2-150 (R257)

FR3-150 (R262)

FR4-150 (R261)

E-4
APPENDIX F—EFFECT OF AILERON DEFLECTION (SUPPLEMENTARY)

(a) Lift coefficient (negative δA) (b) Lift coefficient (positive δA) 

(c) Hinge-moment coefficient (negative δA) (d) Hinge-moment coefficient (positive δA) 

(e) Lift coefficient vs δA (f) Hinge-moment coefficient vs δA

FIGURE F-1. EFFECT OF AILERON DEFLECTION ON LIFT AND HINGE-MOMENT 

COEFFICIENTS; IRT-CS10 (ICING CONDITION 1); Re = 1.8x106

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.60

-0.40

-0.20

0.00

0.20

0.40

0.60

0.80

Lift coefficient

δA (deg)

0 (R018)

-2.5 (R019)

-5 (R020)

-10 (R021)

-15 (R022)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.60

-0.40

-0.20

0.00

0.20

0.40

0.60

0.80

Lift coefficient

δA (deg)

0 (R018)

2.5 (R017)

5 (R016)

10 (R015)

15 (R014)

20 (R013)

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.60

-0.40

-0.20

0.00

0.20

0.40

0.60

0.80

1.00

1.20

1.40

1.60

1.80

Lift coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.20

-0.15

-0.10

-0.05

0.00

0.05

0.10

0.15

Hinge moment coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.25

-0.20

-0.15

-0.10

-0.05

0.00

0.05

Hinge moment coefficient

δA (deg)

0 (R018)

2.5 (R017)

5 (R016)

10 (R015)

15 (R014)

20 (R013)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.25

-0.20

-0.15

-0.10

-0.05

0.00

0.05

Hinge moment coefficient

δA (deg)

0 (R018)

-2.5 (R019)

-5 (R020)

-10 (R021)

-15 (R022)

F-1
(a) Lift coefficient (negative δA) (b) Lift coefficient (positive δA) 

(c) Hinge-moment coefficient (negative δA) (d) Hinge-moment coefficient (positive δA) 

(e) Lift coefficient vs δA (f) Hinge-moment coefficient vs δA

FIGURE F-2. EFFECT OF AILERON DEFLECTION ON LIFT AND HINGE-MOMENT 

COEFFICIENTS; IRT-IS10 (ICING CONDITION 2); Re = 1.8x106

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.60

-0.40

-0.20

0.00

0.20

0.40

0.60

0.80

1.00

1.20

Lift coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.20

-0.15

-0.10

-0.05

0.00

0.05

0.10

0.15

Hinge moment coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.60

-0.40

-0.20

0.00

0.20

0.40

0.60

0.80

Lift coefficient

δA (deg)

0 (R110)

-2.5 (R028)

-5 (R029)

-10 (R030)

-15 (R031)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.60

-0.40

-0.20

0.00

0.20

0.40

0.60

0.80

Lift coefficient

δA (deg)

0 (R110)

2.5 (R104)

5 (R105)

10 (R106)

15 (R107)

20 (R108)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.10

-0.05

0.00

0.05

0.10

0.15

Hinge moment coefficient

δA (deg)

0 (R110)

-2.5 (R028)

-5 (R029)

-10 (R030)

-15 (R031)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.25

-0.20

-0.15

-0.10

-0.05

0.00

0.05

Hinge moment coefficient

δA (deg)

0 (R110)

2.5 (R104)

5 (R105)

10 (R106)

15 (R107)

20 (R108)

F-2
(a) Lift coefficient (negative δA) (b) Lift coefficient (positive δA) 

(c) Hinge-moment coefficient (negative δA) (d) Hinge-moment coefficient (positive δA) 

(e) Lift coefficient vs δA (f) Hinge-moment coefficient vs δA

FIGURE F-3. EFFECT OF AILERON DEFLECTION ON LIFT AND HINGE-MOMENT 

COEFFICIENTS; IRT-SC5 (ICING CONDITION 3); Re = 1.8x106

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.80

-0.60

-0.40

-0.20

0.00

0.20

0.40

0.60

0.80

1.00

Lift coefficient

δA (deg)

0 (R055)

-2.5 (R054)

-5 (R053)

-10 (R052)

-15 (R051)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.80

-0.60

-0.40

-0.20

0.00

0.20

0.40

0.60

0.80

1.00

Lift coefficient

δA (deg)

0 (R055)

2.5 (R058)

5 (R059)

10 (R060)

15 (R061)

20 (R062)

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.80

-0.60

-0.40

-0.20

0.00

0.20

0.40

0.60

0.80

1.00

1.20

1.40

1.60

1.80

Lift coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.25

-0.20

-0.15

-0.10

-0.05

0.00

0.05

0.10

0.15

Hinge moment coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.25

-0.20

-0.15

-0.10

-0.05

0.00

0.05

0.10

Hinge moment coefficient

δA (deg)

0 (R055)

2.5 (R058)

5 (R059)

10 (R060)

15 (R061)

20 (R062)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.10

-0.05

0.00

0.05

0.10

0.15

Hinge moment coefficient

δA (deg)

0 (R055)

-2.5 (R054)

-5 (R053)

-10 (R052)

-15 (R051)

F-3
(a) Lift coefficient (negative δA) (b) Lift coefficient (positive δA) 

(c) Hinge-moment coefficient (negative δA) (d) Hinge-moment coefficient (positive δA) 

(e) Lift coefficient vs δA (f) Hinge-moment coefficient vs δA

FIGURE F-4. EFFECT OF AILERON DEFLECTION ON LIFT AND HINGE-MOMENT 

COEFFICIENTS; IRT-CS2 (ICING CONDITION 4); Re = 1.8x106

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.60

-0.40

-0.20

0.00

0.20

0.40

0.60

0.80

1.00

Lift coefficient

δA (deg)

0 (R070)

2.5 (R071)

5 (R072)

10 (R073)

15 (R074)

20 (R075)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.80

-0.60

-0.40

-0.20

0.00

0.20

0.40

0.60

0.80

Lift coefficient

δA (deg)

0 (R070)

-2.5 (R069)

-5 (R068)

-10 (R067)

-15 (R066)

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.80

-0.60

-0.40

-0.20

0.00

0.20

0.40

0.60

0.80

1.00

1.20

1.40

1.60

Lift coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.25

-0.20

-0.15

-0.10

-0.05

0.00

0.05

0.10

0.15

0.20

Hinge moment coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.10

-0.05

0.00

0.05

0.10

0.15

0.20

Hinge moment coefficient

δA (deg)

0 (R070)

-2.5 (R069)

-5 (R068)

-10 (R067)

-15 (R066)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.25

-0.20

-0.15

-0.10

-0.05

0.00

0.05

0.10

Hinge moment coefficient

δA (deg)

0 (R070)

2.5 (R071)

5 (R072)

10 (R073)

15 (R074)

20 (R075)

F-4
(a) Lift coefficient (negative δA) (b) Lift coefficient (positive δA) 

(c) Hinge-moment coefficient (negative δA) (d) Hinge-moment coefficient (positive δA) 

(e) Lift coefficient vs δA (f) Hinge-moment coefficient vs δA

FIGURE F-5. EFFECT OF AILERON DEFLECTION ON LIFT AND HINGE-MOMENT 

COEFFICIENTS; LS-CS10 (ICING CONDITION 1); Re = 1.8x106

-10 -5 0 5 10 15 20 2

Angle of attack (deg)

5

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

Lift coefficient

δA (deg)

0 (R334)

2.5 (R333)

5 (R335)

10 (R336)

15 (R337)

20 (R338)

-10 -5 0 5 10 15 20 2

Angle of attack (deg)

5

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

Lift coefficient

δA (deg)

0 (R334)

-2.5 (R332)

-5 (R331)

-10 (R330)

-15 (R329)

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.60

-0.40

-0.20

0.00

0.20

0.40

0.60

0.80

1.00

1.20

1.40

Lift coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.20

-0.15

-0.10

-0.05

0.00

0.05

0.10

0.15

Hinge moment coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.25

-0.20

-0.15

-0.10

-0.05

0.00

0.05

Hinge moment coefficient

δA (deg)

0 (R334)

2.5 (R333)

5 (R335)

10 (R336)

15 (R337)

20 (R338)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.10

-0.05

0.00

0.05

0.10

0.15

Hinge moment coefficient

δA (deg)

0 (R334)

-2.5 (R332)

-5 (R331)

-10 (R330)

-15 (R329)

F-5
(a) Lift coefficient (negative δA) (a) Lift coefficient (positive δA) 

(c) Hinge-moment coefficient (negative δA) (d) Hinge-moment coefficient (positive δA) 

(e) Lift coefficient vs δA (f) Hinge-moment coefficient vs δA

FIGURE F-6. EFFECT OF AILERON DEFLECTION ON LIFT AND HINGE-MOMENT 

COEFFICIENTS; LS-IS10 (ICING CONDITION 2); Re = 1.8x106

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

Lift coefficient

δA (deg)

0 (R172)

2.5 (R173)

5 (R174)

10 (R175)

15 (R176)

20 (R177)

-10 -5 0 5 10 15 20 2

Angle of attack (deg)

5

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

Lift coefficient

δA (deg)

0 (R172)

-2.5 (R171)

-5 (R170)

-10 (R169)

-15 (R168)

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.80

-0.60

-0.40

-0.20

0.00

0.20

0.40

0.60

0.80

1.00

1.20

1.40

Lift coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.25

-0.20

-0.15

-0.10

-0.05

0.00

0.05

0.10

0.15

Hinge moment coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.10

-0.05

0.00

0.05

0.10

0.15

Hinge moment coefficient

δA (deg)

0 (R172)

-2.5 (R171)

-5 (R170)

-10 (R169)

-15 (R168)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.25

-0.20

-0.15

-0.10

-0.05

0.00

0.05

0.10

Hinge moment coefficient

δA (deg)

0 (R172)

2.5 (R173)

5 (R174)

10 (R175)

15 (R176)

20 (R177)

F-6
(a) Lift coefficient (negative δA) (b) Lift coefficient (positive δA) 

(c) Hinge-moment coefficient (negative δA) (d) Hinge-moment coefficient (positive δA) 

(e) Lift coefficient vs δA (f) Hinge-moment coefficient vs δA

FIGURE F-7. EFFECT OF AILERON DEFLECTION ON LIFT AND HINGE-MOMENT 

COEFFICIENTS; LS-SC5 (ICING CONDITION 3); Re = 1.8x106

-10 -5 0 5 10 15 20 2

Angle of attack (deg)

5

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

1.2

Lift coefficient

δA (deg)

0 (R187)

2.5 (R188)

5 (R189)

10 (R190)

15 (R191)

20 (R192)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Lift coefficient

δA (deg)

0 (R187)

-2.5 (R185)

-5 (R184)

-10 (R183)

-15 (R182)

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.80

-0.60

-0.40

-0.20

0.00

0.20

0.40

0.60

0.80

1.00

1.20

1.40

1.60

1.80

Lift coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.25

-0.20

-0.15

-0.10

-0.05

0.00

0.05

0.10

0.15

0.20

Hinge moment coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.10

-0.05

0.00

0.05

0.10

0.15

0.20

Hinge moment coefficient

δA (deg)

0 (R187)

-2.5 (R185)

-5 (R184)

-10 (R183)

-15 (R182)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.25

-0.20

-0.15

-0.10

-0.05

0.00

0.05

0.10

Hinge moment coefficient

δA (deg)

0 (R187)

2.5 (R188)

5 (R189)

10 (R190)

15 (R191)

20 (R192)

F-7
(a) Lift coefficient (negative δA) (b) Lift coefficient (positive δA) 

(c) Hinge-moment coefficient (negative δA) (d) Hinge-moment coefficient (positive δA) 

(e) Lift coefficient vs δA (f) Hinge-moment coefficient vs δA

FIGURE F-8. EFFECT OF AILERON DEFLECTION ON LIFT AND HINGE-MOMENT 

COEFFICIENTS; LS-CS2 (ICING CONDITION 4); Re = 1.8x106

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Lift coefficient

δA (deg)

0 (R206)

2.5 (R207)

5 (R208)

10 (R209)

15 (R210)

20 (R211)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

Lift coefficient

δA (deg)

0 (R206)

-2.5 (R205)

-5 (R204)

-10 (R203)

-15 (R202)

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.80

-0.60

-0.40

-0.20

0.00

0.20

0.40

0.60

0.80

1.00

1.20

1.40

1.60

Lift coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.25

-0.20

-0.15

-0.10

-0.05

0.00

0.05

0.10

0.15

0.20

Hinge moment coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.10

-0.05

0.00

0.05

0.10

0.15

0.20

Hinge moment coefficient

δA (deg)

0 (R206)

-2.5 (R205)

-5 (R204)

-10 (R203)

-15 (R202)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.25

-0.20

-0.15

-0.10

-0.05

0.00

0.05

0.10

Hinge moment coefficient

δA (deg)

0 (R206)

2.5 (R207)

5 (R208)

10 (R209)

15 (R210)

20 (R211)

F-8
(a) Lift coefficient (negative δA) (b) Lift coefficient (positive δA) 

(c) Hinge-moment coefficient (negative δA) (d) Hinge-moment coefficient (positive δA) 

(e) Lift coefficient vs δA (f) Hinge-moment coefficient vs δA

FIGURE F-9. EFFECT OF AILERON DEFLECTION ON LIFT AND HINGE-MOMENT 

COEFFICIENTS; LS-CS22N (ICING CONDITION 5); Re = 1.8x106

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.60

-0.40

-0.20

0.00

0.20

0.40

0.60

0.80

1.00

1.20

Lift coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.20

-0.15

-0.10

-0.05

0.00

0.05

0.10

0.15

Hinge moment coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

Lift coefficient

δA (deg)

0 (R414)

-2.5 (R410)

-5 (R411)

-10 (R412)

-15 (R413)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

Lift coefficient

δA (deg)

0 (R414)

2.5 (R409)

5 (R408)

10 (R406)

15 (R405)

20 (R404)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.10

-0.05

0.00

0.05

0.10

0.15

Hinge moment coefficient

δA (deg)

0 (R414)

-2.5 (R410)

-5 (R411)

-10 (R412)

-15 (R413)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.25

-0.20

-0.15

-0.10

-0.05

0.00

0.05

Hinge moment coefficient

δA (deg)

0 (R414)

2.5 (R409)

5 (R408)

10 (R406)

15 (R405)

20 (R404)

F-9
(a) Lift coefficient (negative δA) (b) Lift coefficient (positive δA) 

(c) Hinge-moment coefficient (negative δA) (d) Hinge-moment coefficient (positive δA) 

(e) Lift coefficient vs δA (f) Hinge-moment coefficient vs δA

FIGURE F-10. EFFECT OF AILERON DEFLECTION ON LIFT AND HINGE-MOMENT 

COEFFICIENTS; LS-CS22S (ICING CONDITION 5); Re = 1.8x106

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.60

-0.40

-0.20

0.00

0.20

0.40

0.60

0.80

1.00

1.20

Lift coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.25

-0.20

-0.15

-0.10

-0.05

0.00

0.05

0.10

0.15

Hinge moment coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

Lift coefficient

δA (deg)

0 (R442)

-2.5 (R437)

-5 (R438)

-10 (R440)

-15 (R441)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

Lift coefficient

δA (deg)

0 (R442)

2.5 (R436)

5 (R435)

10 (R434)

15 (R433)

20 (R432)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.10

-0.05

0.00

0.05

0.10

0.15

Hinge moment coefficient

δA (deg)

0 (R442)

-2.5 (R437)

-5 (R438)

-10 (R440)

-15 (R441)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.25

-0.20

-0.15

-0.10

-0.05

0.00

0.05

0.10

Hinge moment coefficient

δA (deg)

0 (R442)

2.5 (R436)

5 (R435)

10 (R434)

15 (R433)

20 (R432)

F-10
(a) Lift coefficient (negative δA) (b) Lift coefficient (positive δA) 

(c) Hinge-moment coefficient (negative δA) (d) Hinge-moment coefficient (positive δA) 

(e) Lift coefficient vs δA (f) Hinge-moment coefficient vs δA

FIGURE F-11. EFFECT OF AILERON DEFLECTION ON LIFT AND HINGE-MOMENT 

COEFFICIENTS; LS-IPSF22 (ICING CONDITION 6); Re = 1.8x106

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

Lift coefficient

δA (deg)

0 (R348)

2.5 (R344)

5 (R345)

10 (R351)

15 (R349)

20 (R350)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

Lift coefficient

δA (deg)

0 (R348)

-2.5 (R343)

-5 (R342)

-10 (R341)

-15 (R340)

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.60

-0.40

-0.20

0.00

0.20

0.40

0.60

0.80

1.00

1.20

1.40

Lift coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.20

-0.15

-0.10

-0.05

0.00

0.05

0.10

0.15

Hinge moment coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.10

-0.05

0.00

0.05

0.10

0.15

Hinge moment coefficient

δA (deg)

0 (R348)

-2.5 (R343)

-5 (R342)

-10 (R341)

-15 (R340)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.3

-0.2

-0.2

-0.1

-0.1

0.0

0.1

Hinge moment coefficient

δA (deg)

0 (R348)

2.5 (R344)

5 (R345)

10 (R351)

15 (R349)

20 (R350)

F-11
(a) Lift coefficient (negative δA) (b) Lift coefficient (positive δA) 

(c) Hinge-moment coefficient (negative δA) (d) Hinge-moment coefficient (positive δA) 

(e) Lift coefficient vs δA (f) Hinge-moment coefficient vs δA

FIGURE F-12. EFFECT OF AILERON DEFLECTION ON LIFT AND HINGE-MOMENT 

COEFFICIENTS; LR-CS10 (ICING CONDITION 1); Re = 1.8x106

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.60

-0.40

-0.20

0.00

0.20

0.40

0.60

0.80

1.00

1.20

1.40

Lift coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.25

-0.20

-0.15

-0.10

-0.05

0.00

0.05

0.10

0.15

Hinge moment coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

Lift coefficient

δA (deg)

0 (R363)

2.5 (R357)

5 (R358)

10 (R359)

15 (R360)

20 (R362)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

Lift coefficient

δA (deg)

0 (R363)

-2.5 (R356)

-5 (R355)

-10 (R354)

-15 (R353)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.10

-0.05

0.00

0.05

0.10

0.15

Hinge moment coefficient

δA (deg)

0 (R363)

-2.5 (R356)

-5 (R355)

-10 (R354)

-15 (R353)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.25

-0.20

-0.15

-0.10

-0.05

0.00

0.05

Hinge moment coefficient

δA (deg)

0 (R363)

2.5 (R357)

5 (R358)

10 (R359)

15 (R360)

20 (R362)

F-12
(a) Lift coefficient (negative δA) (b) Lift coefficient (positive δA) 

(c) Hinge-moment coefficient (negative δA) (d) Hinge-moment coefficient (positive δA) 

(e) Lift coefficient vs δA (f) Hinge-moment coefficient vs δA

FIGURE F-13. EFFECT OF AILERON DEFLECTION ON LIFT AND HINGE-MOMENT 

COEFFICIENTS; LR-IS10 (ICING CONDITION 2); Re = 1.8x106

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.60

-0.40

-0.20

0.00

0.20

0.40

0.60

0.80

1.00

1.20

1.40

Lift coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.25

-0.20

-0.15

-0.10

-0.05

0.00

0.05

0.10

0.15

Hinge moment coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

Lift coefficient

δA (deg)

0 (R240)

-2.5 (R242)

-5 (R241)

-10 (R199)

-15 (R197)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

Lift coefficient

δA (deg)

0 (R240)

2.5 (R243)

5 (R244)

10 (R245)

15 (R246)

20 (R247)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.10

-0.05

0.00

0.05

0.10

0.15

Hinge moment coefficient

δA (deg)

0 (R240)

-2.5 (R242)

-5 (R241)

-10 (R199)

-15 (R197)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.25

-0.20

-0.15

-0.10

-0.05

0.00

0.05

0.10

Hinge moment coefficient

δA (deg)

0 (R240)

2.5 (R243)

5 (R244)

10 (R245)

15 (R246)

20 (R247)

F-13
(a) Lift coefficient (negative δA) (b) Lift coefficient (positive δA) 

(c) Hinge-moment coefficient (negative δA) (d) Hinge-moment coefficient (positive δA) 

(e) Lift coefficient vs δA (f) Hinge-moment coefficient vs δA

FIGURE F-14. EFFECT OF AILERON DEFLECTION ON LIFT AND HINGE-MOMENT 

COEFFICIENTS; LR-SC5 (ICING CONDITION 3); Re = 1.8x106

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.80

-0.60

-0.40

-0.20

0.00

0.20

0.40

0.60

0.80

1.00

1.20

1.40

1.60

1.80

Lift coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.25

-0.20

-0.15

-0.10

-0.05

0.00

0.05

0.10

0.15

0.20

Hinge moment coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Lift coefficient

δA (deg)

0 (R216)

-2.5 (R218)

-5 (R217)

-10 (R215)

-15 (R214)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

1.2

Lift coefficient

δA (deg)

0 (R216)

2.5 (R219)

5 (R220)

10 (R221)

15 (R222)

20 (R223)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.15

-0.10

-0.05

0.00

0.05

0.10

0.15

0.20

Hinge moment coefficient

δA (deg)

0 (R216)

-2.5 (R218)

-5 (R217)

-10 (R215)

-15 (R214)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.25

-0.20

-0.15

-0.10

-0.05

0.00

0.05

0.10

Hinge moment coefficient

δA (deg)

0 (R216)

2.5 (R219)

5 (R220)

10 (R221)

15 (R222)

20 (R223)

F-14
(a) Lift coefficient (negative δA) (b) Lift coefficient (positive δA) 

(c) Hinge-moment coefficient (negative δA) (d) Hinge-moment coefficient (positive δA) 

(e) Lift coefficient vs δA (f) Hinge-moment coefficient vs δA

FIGURE F-15. EFFECT OF AILERON DEFLECTION ON LIFT AND HINGE-MOMENT 

COEFFICIENTS; LR-CS2 (ICING CONDITION 4); Re = 1.8x106

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.80

-0.60

-0.40

-0.20

0.00

0.20

0.40

0.60

0.80

1.00

1.20

1.40

1.60

Lift coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.25

-0.20

-0.15

-0.10

-0.05

0.00

0.05

0.10

0.15

Hinge moment coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

Lift coefficient

δA (deg)

0 (R232)

-2.5 (R230)

-5 (R229)

-10 (R228)

-15 (R227)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

Lift coefficient

δA (deg)

0 (R232)

2.5 (R231)

5 (R234)

10 (R235)

15 (R236)

20 (R237)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.10

-0.05

0.00

0.05

0.10

0.15

Hinge moment coefficient

δA (deg)

0 (R232)

-2.5 (R230)

-5 (R229)

-10 (R228)

-15 (R227)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.25

-0.20

-0.15

-0.10

-0.05

0.00

0.05

0.10

Hinge moment coefficient

δA (deg)

0 (R232)

2.5 (R231)

5 (R234)

10 (R235)

15 (R236)

20 (R237)

F-15
(a) Lift coefficient (negative δA) (b) Lift coefficient (positive δA) 

(c) Hinge-moment coefficient (negative δA) (d) Hinge-moment coefficient (positive δA) 

(e) Lift coefficient vs δA (f) Hinge-moment coefficient vs δA

FIGURE F-16. EFFECT OF AILERON DEFLECTION ON LIFT AND HINGE-MOMENT 

COEFFICIENTS; LR-CS22N (ICING CONDITION 5); Re = 1.8x106

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

Lift coefficient

δA (deg)

0 (R459)

2.5 (R450)

5 (R449)

10 (R448)

15 (R447)

20 (R446)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.4

-0.2

0.0

0.2

0.4

0.6

Lift coefficient

δA (deg)

0 (R459)

-2.5 (R455)

-5 (R456)

-10 (R457)

-15 (R458)

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.40

-0.20

0.00

0.20

0.40

0.60

0.80

1.00

Lift coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.20

-0.15

-0.10

-0.05

0.00

0.05

0.10

0.15

Hinge moment coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.15

-0.10

-0.05

0.00

0.05

0.10

0.15

Hinge moment coefficient

δA (deg)

0 (R459)

-2.5 (R455)

-5 (R456)

-10 (R457)

-15 (R458)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.25

-0.20

-0.15

-0.10

-0.05

0.00

0.05

0.10

Hinge moment coefficient

δA (deg)

0 (R459)

2.5 (R450)

5 (R449)

10 (R448)

15 (R447)

20 (R446)

F-16
(a) Lift coefficient (negative δA) (b) Lift coefficient (positive δA) 

(c) Hinge-moment coefficient (negative δA) (d) Hinge-moment coefficient (positive δA) 

(e) Lift coefficient vs δA (f) Hinge-moment coefficient vs δA

FIGURE F-17. EFFECT OF AILERON DEFLECTION ON LIFT AND HINGE-MOMENT 

COEFFICIENTS; LR-CS22S (ICING CONDITION 5); Re = 1.8x106

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.40

-0.20

0.00

0.20

0.40

0.60

0.80

Lift coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-15 -10 -5 0 5 10 15 20

Aileron Deflection (deg)

-0.20

-0.15

-0.10

-0.05

0.00

0.05

0.10

0.15

Hinge moment coefficient

AOA (deg)

-8

-4

0

4

8

12

16

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.4

-0.2

0.0

0.2

0.4

0.6

Lift coefficient

δA (deg)

0 (R468)

-2.5 (R471)

-5 (R472)

-10 (R473)

-15 (R475)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.4

-0.2

0.0

0.2

0.4

0.6

Lift coefficient

δA (deg)

0 (R468)

2.5 (R470)

5 (R469)

10 (R463)

15 (R462)

20 (R461)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.15

-0.10

-0.05

0.00

0.05

0.10

0.15

Hinge moment coefficient

δA (deg)

0 (R468)

-2.5 (R471)

-5 (R472)

-10 (R473)

-15 (R475)

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.25

-0.20

-0.15

-0.10

-0.05

0.00

0.05

Hinge moment coefficient

δA (deg)

0 (R468)

2.5 (R470)

5 (R469)

10 (R463)

15 (R462)

20 (R461)

F-17/F-18
APPENDIX G—COMPUTATION OF LEWICE ICE SHAPE PROFILES

The ice accretion code used in this study is LEWICE v2.0 [G-1], developed by NASA Glenn 

Research Center. LEWICE is commonly used in the industry to determine ice shape profiles, 

water droplet impingement patterns, water or ice mass flux, and chordwise extent of ice growth 

on the body of interest. Due to the physical modeling and computational methods employed in 

the software, LEWICE executes very rapidly and thus can be used to perform extensive analysis 

as long as the assumptions inherent in the software are not violated. This software uses the clean 

body geometry along with the aerodynamic and meteorological conditions to compute an ice 

shape. The output of the software is a 2-dimensional (2D) ice shape profile that can be 

calculated for several locations along the span of the wing. The approach taken in this work was 

to calculate ice shapes at five locations along the span of the 3-dimensional (3D) wing and then 

construct a full 3D ice shape by lofting between sections. The five locations selected for the ice 

shape computations were 0% (wing root), 15%, 50%, 85%, and 100% (wing tip) semispan. 

In order to use the 2D calculation for construction of the ice shape, adjustments were made to 

account for the 3D flow effects and how these effects impact the droplet trajectory, heat transfer, 

and ice growth calculations. From the work of Dorsch and Brun [G-2], the calculation of droplet 

impingement on a swept wing can be performed by taking a cross-section normal to the leading 

edge and adjusting the velocity by use of the following expression: 

n = VV ∞cosΛ (G-1)

where Vn is the velocity in the plane normal to wing leading edge. 

In addition to this velocity adjustment, α was modified to match the pressure profile at the 

leading-edge region as determined from a 3D Navier-Stokes calculation of pressures over the 

entire wing surface. 

The 3D Navier-Stokes computations were performed at Wichita State University with the clean 

wing. In the computations, the NASA Glenn Icing Research Tunnel (IRT) walls were included 

to simulate the tunnel wall effects on the wing flow field. The α used in the analysis were 4° and 

6°, to match the geometric angles of attack in the icing tests. Analysis pressures for streamwise 

sections at 15%, 50%, and 85% semispan were compared with experimental pressures 

distributions obtained in the IRT facility. Good correlation between experiment and analysis was 

demonstrated. From the computed flow fields, pressure distributions were obtained for the four 

wing sections normal to the wing leading edge (at 15%, 50%, and 85% semispan, and wing tip) 

and for the streamwise section at the wing root. 

The two adjustments described above allow the velocity and pressure profile of the 2D 

calculation to be similar to those of the actual 3D flow. An alternative method to matching the 

pressure profile is to change the angle of attack in the 2D calculations so that the lift of the airfoil 

is equal to the lift of the wing at the cross section of interest or to simply adjust the angle of 

attack geometrically. These approaches were not attempted for this investigation; rather, the 

former method was employed throughout the study. 

G-1 
The exception to this approach was the calculation of the ice shape at the root of the wing. In 

that location, a section cut perpendicular to the leading edge would intersect the floor of the 

tunnel. In addition, the airflow in that region was affected by the presence of the tunnel floor. 

Flow visualization suggests that the flow was parallel to the tunnel walls. As such, the LEWICE 

calculation for the root section was performed using the geometry corresponding to a crosssectional cut parallel to the tunnel walls. 

The temperature input for the LEWICE software was also adjusted to approximate the conditions 

present on the actual wing model. LEWICE requires the static temperature of the free stream as 

input. Since the input velocity was reduced, as indicated in equation G-1, the resulting total 

temperature profile for the airfoil would be reduced by an amount approximately equal to 

p

n

C

VV

2

22

∞ − (G-2)

where Cp is the specific heat of air. 

Thus, the input free-stream static temperature was modified such that the free-stream total 

temperature matched that of the 3D flow. The expression used to make this adjustment is 

provided in equation G-3. 

+= Λ ∞ 2

2

sin

2 p

f,sn,s C

V TT (G-3)

where Ts,n and Ts,f are the static temperature in the plane normal to wing leading edge and freestream plane, respectively. 

In summary, the aerodynamic and icing conditions input for LEWICE analyses to generate the 

seven LEWICE ice shapes are documented in table G-1. 

REFERENCES 

G-1. Wright, W.B., “User Manual for the NASA Glen Ice Accretion Code LEWICE (Version 

2.0),” NASA CR-1999-2094009, September 1999. 

G-2. Dorsch, R.G. and Brun, R.J., “A Method for Determining Cloud-Droplet Impingement on 

Swept Wings,” NACA TN 2931, April 1953. 

G-2 
TABLE G-1. FLOW AND ICING CONDITIONS INPUT FOR LEWICE ICE SHAPES 

Ice Shape 

LWC 

(g/m3

) 

MVD 

(μm) 

τ

(min) 

RH 

(%) 

Pstatic

(N/m2) 

Tstatic

(K) 

Spanwise 

Section 

C 

(m) 

V 

(m/s) 

α

(deg) 

Root 0.640 111.76 4.0 

15% 0.572 98.678 2.8 

50% 0.440 98.678 2.2 

85% 0.308 98.678 0.6 

LS-CS10 0.68 20.0 10.0 100 100000 263 

Tip 0.257 98.678 0.6 

Root 0.640 67.056 4.0 

15% 0.572 59.207 2.8 

50% 0.440 59.207 2.2 

85% 0.308 59.207 0.6 

LS-IS10 0.65 20.0 10.0 100 100000 263 

Tip 0.257 59.207 0.6 

Root 0.640 89.989 4.0 

15% 0.572 79.456 2.8 

50% 0.440 79.456 2.2 

85% 0.308 79.456 0.6 

LS-SC5 0.51 14.5 5.0 100 100000 263 

Tip 0.257 79.456 0.6 

Root 0.640 111.76 4.0 

15% 0.572 98.678 2.8 

50% 0.440 98.678 2.2 

85% 0.308 98.678 0.6 

LS-CS2 0.68 20.0 2.0 100 100000 263 

Tip 0.257 98.678 0.6 

Root 0.640 111.76 4.0 

15% 0.572 98.678 2.8 

50% 0.440 98.678 2.2 

85% 0.308 98.678 0.6 

LS-CS22N 0.68 20.0 22.5 100 100000 263 

Tip 0.257 98.678 0.6 

Root 0.640 111.76 4.0 

15% 0.580 111.76 2.8 

50% 0.448 111.76 2.2 

85% 0.314 111.76 0.6 

LS-CS22S 0.68 20.0 22.5 100 100000 263 

Tip 0.257 111.76 0.6 

Root 0.640 67.056 4.0 

15% 0.572 59.207 2.8 

50% 0.440 59.207 2.2 

85% 0.308 59.207 0.6 

LS-IPSF22 0.46 20.0 22.5 100 100000 263 

Tip 0.257 59.207 0.6 

G-3/G-4 
APPENDIX H—COMPARISON OF TWO- AND THREE-DIMENSIONAL WING ICE

SHAPE TRACES AND LIFT DATA 

H.1 COMPARISON OF TWO- AND THREE-DIMENSIONAL ICE SHAPES. 

The icing conditions for ice shape IRT-SC5, listed in tables 2-1 and 2-2 of this report, were 

obtained by scaling the conditions used in an ice accretion test conducted with a two-dimensional 

(2D) 36-inch chord GLC-305 airfoil [H-1]. 

The scale conditions were determined using the Ruff method [H-2] for sea level testing with 

constant velocity. This method was developed for scaling 2D models to permit simulating a 

larger model by testing a smaller model, or to permit testing at one set of test conditions to 

simulate conditions that were not attainable in the facility being used. For either simulation, the 

objective is to produce a scaled ice shape that simulates the reference shape; in fact, when 

normalized by the model chord, the two shapes should match in overall size and in locations and 

size of individual features. The Ruff method matches scale and reference values of those 

nondimensional similarity parameters that have been demonstrated to have the strongest 

influence on the ice accretion quantity and shape. The dimensionless similarity parameters 

include the modified inertia parameter, K0, which relates to the droplet trajectories; the 

accumulation parameter, Ac, which is a measure of the quantity of ice that can potentially accrete 

on the model; and the freezing fraction, n0, which is the proportion of water impinging at 

stagnation that actually freezes. Another energy parameter that can be used is the water energy 

transfer parameter, φ. This parameter has units of temperature and collects all the terms in the 

energy balance that relate to energy carried to the surface by water drops. For rime ice, it is 

convenient and acceptable to set the scale velocity equal to the reference, although other methods 

of finding scale velocity have been shown to give better results for glaze ice. The five equations 

formed by matching the scale and reference values of the three nondimensional similarity 

parameters, plus the water energy parameter and the velocity can be solved to find the scale test 

conditions. 

The method outlined in the previous paragraph has only been validated to scale 2D (straight 

wing) models. There have been no studies of 2D to three-dimensional (3D) (swept wing) 

scaling, and a crucial question is how valid are the 2D parameters, some of which apply only to 

the stagnation line, might be for swept wings. However, scaling between two geometrically 

similar swept wing models of different size was studied by ONERA [H-3 and H-4] by applying 

most of the same similarity parameters that Ruff later used. Although direct comparisons of 

scale and reference ice shapes were not possible, photographs suggested that quantities and 

features of the reference accretions were simulated well by the scale tests. These results 

provided some encouragement that the Ruff method might be valid for 2D to 3D scaling. To test 

this possibility in a very limited way, the present study included rime tests with a 2D model as a 

reference and compared those ice shapes with ones from a 3D model at scaled conditions. 

The strategy was to apply the Ruff scaling calculations outlined above using rime tests with a 36-

in chord, 2D GLC-305 model as the reference. The scale model was to have the same airfoil 

form as the reference but with a chord of 17.4 inches. Rime ice, even for swept wings, has a 

fairly simple shape, lacking the 3D scallops of swept wing glaze. Therefore, the assumption was 

that rime ice accretions would be no different whether on a 17.4-in chord 2D airfoil or on a 3D 

H-1 
airfoil with a mean aerodynamic chord (MAC) of 17.4 inches, at least at the MAC. The leading 

edge radius for the GLC-305 airfoils was taken to be 0.0134c, where c is the chord. The 

reference conditions and resulting scale conditions are shown in the following table: 

c 

(in) 

V 

(mph) 

α

(deg) 

Ttotal

(°F) 

MVD 

(μm) 

LWC 

(g/m3

) 

Time 

(minutes) 

β0

(%) β0Ac

Reference 36.0 201.3 6 11.7 20.0 0.40 16.7 66.3 1.065 

Scaled 17.4 201.3 6 11.7 12.8 0.58 5.6 66.3 1.065 

The relative thickness of ice accreted at the leading edge is proportional to the product of three 

parameters: the collection efficiency at stagnation, β0, the accumulation parameter, Ac, and the 

stagnation freezing fraction, n0. Thus, for rime ice (n0 of 1), the product β0Ac gives a measure of 

the relative size of the ice accretions. Note that for proper scaling, this product must match 

between scale and reference. Other similarity parameters are not important for rime scaling. 

Some adjustments to the calculated scale conditions were necessary. First, because the scale 

drop size determined by the Ruff method was smaller than the minimum value for which the 

icing research tunnel (IRT) cloud was calibrated (14.5 μm), the minimum median volumetric 

diameter (MVD) for the IRT had to be substituted. For a straight wing, this increase in drop size 

would result in a stagnation collection efficiency about 6% higher than the reference. For the 

swept wing, the assumption was made that the change would have the same proportional effect 

as that for the straight wing. Thus, the second adjustment was to reduce the spray time from 5.6 

to 5 minutes to compensate for the increased collection efficiency. Then the revised scale 

conditions were: 

c 

(in) 

V 

(mph) 

α

(deg) 

Ttotal

(°F) 

MVD 

(μm) 

LWC 

(g/m3

) 

Time 

(minutes) 

β0

(%) β0Ac

Reference 36.0 201.3 6 11.7 20.0 0.40 16.7 66.3 1.065 

Scaled 17.4 201.3 6 11.7 14.5 0.58 5.0 70.2 1.003 

Although the scaling calculations were made using a chord of 17.4 inches, the icing tests were 

actually performed with a swept GLC-305 airfoil with a MAC of 18.72 inches. Both the 

collection efficiency and accumulation parameter are dependent on chord. The local chord at 

which icing traces were made also varied with spanwise location; therefore, the scale test 

conditions shown above were valid only for one location. In addition to the change in MAC, the 

liquid water content (LWC) actually tested was not the value desired due to an error in the IRT 

calibration discovered after the completion of tests. Thus, while the spray bar conditions were 

set to give an LWC of 0.58 g/m3

, as recommended from the scale calculations, the actual value is 

believed to have been 0.51 g/m3

. 

Figure H-1 compares a trace of the ice shape obtained from the icing tests with the 2D GLC-305 

with three traces of the ice shape obtained from the 3D swept wing icing tests at the NASA 

Glenn IRT facility. The traces of the 3D ice shape were obtained at wing stations A, B, and C, as 

shown in figure 2-5 of this report. The ice traces at stations A and B were taken normal to the 

H-2 
swept wing leading edge, while the trace at station C was taken in the streamwise direction. As 

discussed in the main body of this report, the swept wing had a GLC-305 airfoil section that was 

in the streamwise direction. Thus, sections taken normal to the leading edge of the swept wing 

were slightly thicker than the streamwise GLC-305 section and had a chord length that was 

approximately 99% of the streamwise airfoil chord length. The airfoil and ice traces presented in 

figure H-1 have been normalized with the normal or streamwise chord of each airfoil section. 

These nondimensional tracings show good agreement between the 2D trace and that of cut A of 

the swept wing, but the relative quantity of ice for cuts B and C are far less than that for the 2D 

trace. The following table shows the relevant similarity parameters for the conditions actually 

tested at the MAC and at each of the tracing locations: 

FIGURE H-1. COMPARISON OF THE ICE SHAPES FOR THE 2D AND 3D MODELS 

-0.05 0.00 0.05 0.10 0.15 0.20 0.25 0.30

X/C

-0.15

-0.10

-0.05

0.00

0.05

0.10

0.15

Y/C

2D Tests

3D Tests (Cut A, refer to figure 2.5)

3D Tests (Cut B, refer to figure 2.5)

3D Tests (Cut C, refer to figure 2.5)

c 

(in) 

V 

(mph) 

α

(deg) 

Ttotal

(°F) 

MVD 

(μm) 

LWC 

(g/m3

) 

Time 

(minutes) 

β0

(%) β0Ac

Reference 36.00 201.3 6 11.7 20.0 0.40 16.7 66.3 1.065 

Scaled (MAC) 18.72 201.3 6 11.7 14.5 0.51 5.0 68.9 0.811 

Scaled (Cut A) 13.80 201.3 6 11.7 14.5 0.51 5.0 74.1 1.182 

Scaled (Cut B) 19.30 201.3 6 11.7 14.5 0.51 5.0 68.3 0.781 

Scaled (Cut C) 25.10 201.3 6 11.7 14.5 0.51 5.0 63.2 0.555 

Note that the scale value of β0Ac is only close to the reference at cut A, where the scale and 

reference are within 11%. The parameter Ac is directly proportional to LWC, and the uncertainty 

in the LWC calibration for the IRT is generally quoted as ±10%, so these numbers are 

reasonably close. The good agreement of the ice tracings for the 2D and 3D at cut A is, 

therefore, consistent with the β0Ac value there, and this result shows that for this rime case, at 

least, it is possible to scale effectively from 2D to 3D. The much smaller scale shapes at cuts B 

and C are predicted by the significantly lower β0Ac values at those locations. 

H-3 
Although the results for nearly matched β0Ac were encouraging, the tests were very limited in 

scope, including only rime ice at one set of reference conditions. For glaze ice, the scallop 

formation on swept wings of large sweep angles makes a direct comparison with 2D ice shapes 

impossible. However, additional tests for glaze ice at several freezing fractions are encouraged 

to determine if useful similarities in 2D and 3D shapes can be identified. For future tests, it 

should also be recognized that the matching of reference and scale similarity parameters applies 

only at one swept wing spanwise location. 

H.2 COMPARISON OF 2D AND 3D CLEAN AND ICED WING AERODYNAMIC

PERFORMANCE.

In this section, the aerodynamic lift performance of the clean and iced GLC-305 2D wing model 

presented in reference H-1 is compared with the lift performance of the clean and iced GLC-305 

3D wing presented in the main body of this report. The 2D and 3D iced wing configurations 

considered below are the ones presented in figure H-1. The aerodynamic performance 

comparisons will be limited to linear lift slope and maximum lift coefficient. 

In comparing the aerodynamic performance of the two wing models, the following differences in 

model geometry and flow-field behavior should be considered: 

• Model size and geometric features. The 2D model was a straight infinite span planar 

wing with a chord of 36 inches. The 3D model was a finite span swept wing with 

geometric twist and taper, and with a MAC of 18.72 inches. 

• The lift data for the 2D model were obtained at Reynolds numbers in the range of 3 to 7.5 

million compared to the 1.8 million Reynolds number used in the 3D wing tests. 

• The flow field about a finite swept wing is three-dimensional and is, therefore, inherently 

different from 2D flow about an infinite straight wing. Three-dimensional effects include 

spanwise flow, leading edge and wing tip vortices, downwash effects, and complex 

separation patterns at high angles of attack as the wing approaches stall. As a result, the 

stall behavior of the 2D infinite span wing will in general be different from that of the 

finite span wing with the same airfoil section. 

• Aerodynamic performance of finite swept wings depends on aspect ratio (AR), taper ratio 

(λ), sweep angle (Λ), Reynolds number (Re), wing twist, and potential fuselage/wing 

interaction effects. Note that the 3D wing was tested with a fuselage like body, as shown 

in figure 2-18 of this report. 

From experimental work and classical aerodynamic theory, a number of simple equations have 

been developed for relating the lift slope of a straight finite wing to that of a 2D wing with the 

same airfoil section. For planar untapered swept wings, basic swept wing theory can be applied 

to correct the lift slope of straight wings for the effects of wing sweep as discussed in 

reference H-5. Sweep corrections account for the fact that, for swept wings only the velocity 

component normal to the wing leading edge is responsible for wing lift and surface pressures. 

The velocity component tangential to the wing leading edge is important only for the 

H-4 
determination of the frictional stresses on the surface. Corrections for sweep effects depend to 

some extent on the method used to convert a straight wing into a swept wing. In general, sweep 

is introduced by rotating or by shearing a straight wing. If rotation is applied to sweep the wing 

(bent-back or yawed wing), then the airfoil section of the unswept wing is the same as the airfoil 

section normal to the leading edge of the swept wing. A sheared wing is obtained by shearing 

backward (or forward) every section of the unswept airfoil, leaving its shape and lateral position 

unchanged. The swept finite GLC-305 wing was obtained by applying the shear method. Thus, 

the airfoil section of the straight 2D wing and the streamwise section of the swept wing were the 

same. However, the airfoil section normal to the swept wing leading edge was slightly thicker 

than that of the standard GLC 305 profile. Note that the simple sweep theory requires the use of 

the lift data for the airfoil section normal to the wing leading edge. A 2D viscous flow analysis 

was conducted with the XFOIL code [H-6] to assess the difference in lift performance between 

the GLC-305 airfoil and the thicker airfoil section normal to the wing leading edge. The results 

showed that the airfoil normal to the wing leading edge had a slightly higher maximum lift 

coefficient but in general the lift performance of both airfoils was very similar over the linear lift 

range. Thus, it was decided to use the baseline GLC-305 lift data in all calculations presented 

below. 

Experimental lift curves for the 2D and 3D clean wing are compared in figure H-2. The main 

differences between the 2D and 3D wing lift curves include the following: 

• Considerably higher lift slope for the 2D wing (i.e., higher lift at a given angle of attack). 

• Angle of zero lift is negative for the 2D wing but slightly positive for the 3D wing. 

• Maximum lift is considerably greater for the 2D wing. 

• Stall behavior for the 2D and 3D wings is considerably different. A gradual stall is 

observed to take place for the 3D wing compared to a sharp stall for the 2D wing. 

FIGURE H-2. COMPARISON OF LIFT CURVES FOR THE CLEAN 2D AND 3D WINGS 

-10 -5 0 5 10 15 20 25

Angle of attack (deg)

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

1.2

Lift coefficient

2D Exp data (Re=3.0x106, M=0.120)

3D Exp data (Re=1.8x106, M=0.185)

Predicted CL using equation H1

H-5 
In the following, the lift characteristics of the clean and iced GLC-305 finite wing will be 

estimated from the lift properties of the clean and iced 2D wing using simple equations 

developed for swept wings. It should be emphasized that stall behavior cannot be predicted by 

these simple schemes. 

The method used to estimate the finite swept wing lift data from that of the 2D wing was as 

follows: 

• Apply standard sweep, aspect ratio, and compressibility corrections to estimate the 3D 

wing lift slope. 

• Estimate effective twist angle for the 3D wing and use this value to correct the 3D wing 

geometric angle of attack. 

• Estimate the effect of Reynolds number on lift slope using experimental data and a 2D 

viscous analysis panel code. 

• Apply a simple formula to estimate maximum lift for the 3D wing from that of the 2D 

wing. 

Due to downwash and geometric twist effects, the lift curve slope of the 3D swept wing was 

reduced compared to that of the 2D wing model. Knowing the 2D lift slope (a0) and the angle of 

attack for zero lift (αL=0), one can predict the lift curve of a finite swept wing from equation H-1 

[H-7] which takes sweep angle and aspect ratio into consideration. Note that all lift slopes in 

equation H-1 are per radian. From reference 24, the 2D slopes (a0) of the clean and iced wing 

were 0.0973 and 0.0977 per degree, respectively. Corresponding lift slope values per radian 

were 5.574 and 5.598. From equation H-1, the lift slopes (CLα) of the clean and iced finite swept 

wings were calculated to be 4.114 and 4.127 per radian or 0.0718 and 0.072 per degree, 

respectively. 

( )2

2

0

2

21

0

0

∞ ⎟ ⋅− ⎠

⎞ ⎜

⎝

⎛ + ⎟

⎟

⎠

⎞

⎜

⎜

⎝

⎛

Λ ⎟ +

⎠

⎞ ⎜

⎝

⎛

⋅ =

MAR a

cos

a AR

ARa CL

π π

α (H-1) 

where AR = 6.8, Λ1/2 = 22° and M∞ = 0.185 for the swept finite wing tested. 

Typically, planar wings with infinite and finite span and identical airfoil sections have the same 

angle of zero lift (αL=0). The GLC-305 swept wing model, however, had a washout of 4° at the 

wing tip, which reduced the effective angle of attack of the 3D wing, thus (αL=0)3D and (αL=0)2D

were not the same. For a finite wing with washout, (αL=0)3D should be the difference of (αL=0)2D

and θeff, where θeff is the effective angle of twist of the whole wing. Note that at zero lift, the 

downwash angle generated by a finite wing is practically zero, thus only wing geometric twist 

causes the 2D and 3D geometric angles of attack to differ. To estimate the effective twist angle 

of the finite swept wing, the spanwise lift distribution of the 3D wing must be considered. Using 

H-6 
the spanwise lift distribution, the spanwise location, where the wing local lift times the span of 

the wing is approximately equal to the total lift of the wing (integral under the curve), is 

determined. From the limited surface pressure data obtained with the 3D wing, an approximate 

spanwise lift distribution was computed and the data was used to determine the spanwise 

location corresponding to the effective twist angle. This location was close to the 50% semispan 

for the angles of attack presented in figure H-3. The geometric twist at the 50% semispan station 

(θ50) was -1.14°. Thus, the effective twist of the wing at low angles of attack was -1.14°. 

Therefore, the difference between the 2D and 3D geometric angles of attack at zero lift was 1.14°

(the 3D angle of attack was 1.14° higher than the 2D geometric angle of attack). 

0 5 10 15 20 25 30 35 40 45 50 55 60

Spanwise station (in)

-0.4

-0.3

-0.2

-0.1

0.0

0.1

0.2

0.3

0.4

0.5

0.6

Section lift coefficient

Wing CL α=-3.23°

α=-1.09°

α=-0.02°

α=2.12°

α=4.30°

α=7.50°

FIGURE H-3. SPANWISE LIFT DISTRIBUTIONS OF THE CLEAN SWEPT WING 

The effect of Reynolds number on the lift behavior is demonstrated in figure H-4 using the 

experimental data of reference H-1 and computational data obtained with the XFOIL code [H-6]. 

Both the experimental and analysis data indicate a significant increase in maximum lift and angle 

of stall as expected. In addition, a small increase in lift slope is observed as Reynolds number is 

increased. Since the airfoil section for the 3D was constant from root to tip, the Re effects for the 

3D wing should be similar to that observed with the 2D airfoil. 

The experimental lift curves of the clean and iced wings are compared in figures H-2 and H-5 

respectively. Each figure provides the 2D and 3D experimental lift curves and the linear part of 

the lift curve for the 3D wing (line with diamond symbols) that was computed from simple wing 

sweep theory. The results indicate that the estimated 3D wing slope is in good agreement with 

the experimental data of the 3D wing. 

Next, an estimate of the maximum lift coefficient for the 3D wing is computed using simple 

wing sweep theory. According to Hoerner [H-5], CLmax for a swept wing can be predicted from 

the maximum lift coefficient (Clmax,2D) of a 2D wing using equation H-2. Again, the formula 

provided is strictly applicable to planar wings and assumes that the airfoil section, in this case 

GLC-305, and the flow are normal to the wing leading edge. 

4

2

2 cos max ClCL Dmax, Λ⋅= /c (H-2)

H-7 
where Λc/4 is the quarter-chord sweep angle, which was 25° for the 3D GLC-305 swept wing. 

Experimental data presented in reference H-5 indicate that equation H-2 underestimates Clmax for 

sweep angles greater than about 20°. 

FIGURE H-4. EFFECT OF REYNOLDS 

NUMBER ON GLC-305 LIFT SLOPE 

FIGURE H-5. COMPARISON OF LIFT 

CURVES FOR THE ICED 2D AND 

3D WINGS 

Lift slopes, maximum lift coefficients, and angles of stall for the 2D and 3D clean and iced wing 

models are summarized in table H-1. The stall angles provided in table H-1 were obtained from 

the experimental data. Due to spanwise flow, leading edge and tip vortex structures, washout, 

and downwash effects, the stall mechanism of finite swept wings is considerably different from 

that of straight 2D wings and in general the angle of stall for 3D wings is higher. 

TABLE H-1. SUMMARY OF LIFT CURVE SLOPES AND CLmax 

Clean Wing Iced Wing 

2D Exp 0.0973 0.0977 

Exp. 0.0765 0.0775 Lift curve slope 

(per degree) 3D Calc.* 0.0718 0.0720 

2D Exp 1.0850 0.9702 

CLmax Exp. 0.8738 0.8955 3D Calc.* 0.8912 0.7969 

2D Exp 10.54 13.23 

αstall (deg) 3D Exp 13.78 15.80 

* Estimates for 3D wing obtained from the 2D experimental data of reference H-1 and simple 

sweep wing theory. 

-6 -4 -2 0 2 4 6 8 10 12 14 1

Angle of attack (deg)

6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

1.6

Lift coefficient

Re=1.8x106, M=0.185 (XFOIL)

Re=3.0x106, M=0.12 (XFOIL)

Re=7.5x106, M=0.21 (XFOIL)

Re=7.5x106, M=0.21 (Exp)

Re=3.0x106, M=0.12 (Exp)

-10 -5 0 5 10 15 20 2

Angle of attack (deg)

5

-0.8

-0.6

-0.4

-0.2

0.0

0.2

0.4

0.6

0.8

1.0

Lift coefficient

2D Exp data (Re=3.5x106, M=0.210)

3D Exp data (Re=1.8x106, M=0.185)

Predicted CL using equation H1

H-8 
H.3 RELATIONSHIP BETWEEN 2D AND 3D WING GEOMETRIC ANGLE OF ATTACK. 

To estimate the effective angle of attack of the 3D wing, the geometric twist and downwash 

angle effects need to be estimated and the geometric angle of attack must be adjusted for these 

effects. These corrections will provide an equivalent geometric angle of attack for the 2D wing. 

Note that for the same lift, the geometric angle of attack of a finite wing is greater than that of an 

infinite span wing. According to Hoerner [H-5], the characteristic of pressure distributions in the 

vicinity of 50% semi-span of a swept wing is the same as that found in conventional infinite span 

straight wing sections. Therefore, by matching pressure distributions at 50% semispan of the 

clean 3D wing model to the coefficient pressure distributions of the clean 2D wing model, the 

relationship between the geometric angles of attack of the 2D and 3D wing models can be 

estimated. Note that this method assumes the Reynolds numbers of the 3D and 2D flows are the 

same, which is not the case here. However, as it was shown in figure H-4, the effect of Re over 

the linear portion of the lift curve was small. The results obtained by applying the pressure 

matching method are presented in figure H-6. By fitting a least-squares straight line to the data 

presented in figure H-2, the relationship between the 2D and 3D geometric angles of attack can 

be obtained. Note that the difference (α2D - α3D) provides an estimate of the average geometric 

twist and downwash angle at 50% semispan. 

FIGURE H-6. RELATIONSHIP OF GEOMETRIC ANGLES OF ATTACK OF THE TWODIMENSIONAL AND THREE-DIMENSIONAL CLEAN WING MODELS 

-6 -4 -2 0 2 4 6 8 1

α of swept wing, α3D (deg)

0

-6

-4

-2

0

2

4

6

8

α of straight wing, 

α2D (deg)

H-9 
H.4 REFERENCES. 

H-1. Addy, H.E., Broeren, A.P., Zoeckler, J.G., and Lee, S., “A Wind Tunnel Study of Icing 

Effects on a Business Jet Airfoil,” AIAA Paper 2003-727, January 2003. 

H-2. Ruff, G.A., “Analysis and Verification of the Icing Scaling Equations,” AEDC-TR-85-

30, Vol. 1 (Rev), March 1986. 

H-3. Armand, C., Charpin, F., Fasso, G., and LeClerc, G., “Techniques and Facilities Used at 

the ONERA Modane Centre for Icing Tests,” AGARD-AR-127, Appendix A6, 

November 1978.

H-4. Charpin, F. and Fasso, G., “Icing Testing in the Large Modane Wind-Tunnel on FullScale and Reduced Scale Models,” NASA TM-75373, March 1979.

H-5. Hoerner, I.S.F. and Borst, H.V., “Fluid Dynamic – Lift,” Hoerner Fluid Dynamics, P. O. 

Box 342, Brick Town, NJ, 1975. 

H-6. Drela, M., “XFOIL 6.6 User Primer,” MIT Aero and Astro Engineering, March 1996. 

H-7. McCormick, B. W., “Aerodynamics Aeronautics and Flight Mechanics,” 2nd edition, John 

Wiley and Sons, 1995. 

H-10