Andy P. Broeren

Glenn Research Center, Cleveland, Ohio

James T. Riley

Federal Aviation Administration, William J. Hughes Technical Center, Atlantic City Airport, New Jersey 

Review of the Aerodynamic Acceptance Test and 

Application to Anti-Icing Fluids Testing in the 

NRC Propulsion and Icing Wind Tunnel

NASA/TM—2012-216014 DOT/FAA/TC–12/32

August 2012
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Andy P. Broeren

Glenn Research Center, Cleveland, Ohio

James T. Riley

Federal Aviation Administration, William J. Hughes Technical Center, Atlantic City Airport, New Jersey 

Review of the Aerodynamic Acceptance Test and 

Application to Anti-Icing Fluids Testing in the 

NRC Propulsion and Icing Wind Tunnel

NASA/TM—2012-216014 DOT/FAA/TC–12/32

August 2012

National Aeronautics and

Space Administration

Glenn Research Center 

Cleveland, Ohio 44135
Acknowledgments

The authors wish to acknowledge the contributions to this report from other members of the research team. Warren Underwood of 

the FAA and John D’Avirro, Michele Pineau and Marco Ruggi of APS Aviation provided invaluable insights into the history and 

characteristics of anti-icing fluids testing and results. The authors also appreciate the contributions of Catherine Clark and Marc 

MacMaster at NRC Canada in understanding and interpreting the wind-tunnel experimental methods and aerodynamic data. The 

support of these individuals is gratefully acknowledged. The NASA Aviation Safety Program, Atmospheric Safety Technologies 

Project, and the FAA Ground Icing Program also support the preparation of this report. 

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Level of Review: This material has been technically reviewed by technical management. 
NASA/TM—2012-216014 1 

Review of the Aerodynamic Acceptance Test and 

Application to Anti-Icing Fluids Testing in the 

NRC Propulsion and Icing Wind Tunnel

Andy P. Broeren

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135 

James T. Riley

Federal Aviation Administration 

William J. Hughes Technical Center

Atlantic City Airport, New Jersey 08405 

Abstract

In recent years, the FAA has worked with Transport Canada, National Research Council of Canada 

(NRC) and APS Aviation, Inc. to develop allowance times for aircraft operations in ice-pellet 

precipitation. These allowance times are critical to ensure safety and efficient operation of commercial

and cargo flights. Wind-tunnel testing with uncontaminated anti-icing fluids and fluids contaminated with 

simulated ice-pellets had been carried out at the NRC Propulsion and Icing Wind Tunnel (PIWT) to 

better understand the flowoff characteristics and resulting aerodynamic effects. The percent lift loss on 

the thin, high-performance wing model tested in the PIWT was determined at 8° angle of attack and used 

as one of the evaluation criteria in determining the allowance times. Because it was unclear as to how 

performance degradations measured on this model were relevant to an actual airplane configuration, some 

means of interpreting the wing model lift loss was deemed necessary. In this report, the lift loss was

related to the loss in maximum lift of a Boeing 737-200ADV airplane through the Aerodynamic 

Acceptance Test (AAT) performed for fluids qualification. This report provides a review of the research 

basis of the AAT in order to understand how this correlation was applied. A loss in maximum lift 

coefficient of 5.24 percent on the B737-200ADV airplane (which was adopted as the threshold in the 

AAT) corresponds to a lift loss of 7.3 percent on the PIWT model at 8° angle of attack. There is 

significant scatter in the data used to develop the correlation related to varying effects of the various antiicing fluids that were tested and other factors. A statistical analysis indicated the upper limit of lift loss on 

the PIWT model was 9.2 percent. Therefore, for cases resulting in PIWT model lift loss from 7.3 to 9.2 

percent, extra scrutiny of the visual observations is required in evaluating fluid performance with 

contamination. Additional research may result in future changes to this correlation. 

Nomenclature

Cl sectional lift coefficient

CL airplane lift coefficient

Cl,max maximum sectional lift coefficient

CL,max maximum airplane lift coefficient

V2 takeoff safety speed

Vs1g 1g stall speed

αB body angle of attack 

αw wing section angle of attack

δ* boundary-layer displacement thickness

δ*fluid boundary-layer displacement thickness of fluid in AAT at a given temperature
NASA/TM—2012-216014 2 

δ*dry boundary-layer displacement thickness of dry flat plate in AAT

δ*l @ –20 °C boundary-layer displacement thickness limit define for AAT at –20 °C 

Acronyms 

AAT Aerodynamic Acceptance Test

AEA Association of European Airlines 

AMIL Anti-Icing Materials International Laboratory

APS APS Aviation, Inc. 

EG ethylene glycol

FAA Federal Aviation Administration 

HOT holdover time

IRT Icing Research Tunnel

KEAS equivalent airspeed, kn

NASA National Aeronautics and Space Administration

NRC National Research Council (Canada)

OAT outside air temperature

PG propylene glycol 

PIWT Propulsion and Icing Wind Tunnel 

POI Principal Operations Inspector 

SEE standard error of the estimate

1.0 Introduction

During the period from 1992 to 2004, many air carriers were allowed to operate in ice-pellet

conditions provided that the flight crew did a contamination check of the wings within 5 min of takeoff

although no holdover times (HOTs) existed for these conditions. This check was conducted from inside 

the aircraft in the case of most passenger aircraft, and from the outside for freighters (which had no 

windows in the cabin). This procedure for operation in ice-pellet conditions was incorporated into the 

deicing and anti-icing programs of many individual air carriers and approved annually by the FAA 

Principal Operations Inspector (POI) for those carriers. 

In 2005, the FAA became concerned about operations in ice-pellet conditions after learning of 

questions raised relating to regional carriers operating in ice pellets with non-slatted aircraft and with less 

experienced flight crews than the major carriers. As a result of these concerns, the FAA issued three 

notices effectively disallowing operations in ice pellets in October 2005 (Refs. 1, 2, and 3). In the winter 

of 2005-2006, this action had a major impact at several airports during ice-pellet events. Due in part to 

concerns raised by these events, the FAA undertook a thorough reassessment of its policy with respect to 

operation in ice-pellet conditions. 

HOTs are determined based on the results of endurance time testing of anti-icing fluids in natural or 

simulated icing conditions. Fluid failure occurs when the fluid is no longer able to effectively absorb the 

incoming precipitation. The endurance times are determined by applying failure criteria using visual cues 

for fluid failure, which vary depending on the icing conditions (snow, freezing drizzle, freezing rain). 

Environmental chamber testing showed that ice pellets generally remain in the frozen state imbedded 

in Type IV anti-icing fluid, and are not absorbed by the fluid in the same manner as other forms of 

precipitation. Using the guidelines for determining anti-icing fluid failure, the presence of a contaminant 

not absorbed by the fluid (remaining imbedded) would be an indication that the fluid has failed. However, 

the imbedded ice pellets were generally not readily detectable by the human eye except through short 

range inspection. This indicated that a contamination check of the wings within 5 min of takeoff in 

operational conditions would be of little value. The difficulty in establishing visual cues for detecting 

fluid failure in ice-pellet conditions has thus far precluded the establishment of practical failure criteria 
NASA/TM—2012-216014 3 

for endurance time testing. Without endurance time testing, HOTs cannot be established and so the FAA 

has never provided HOTs for operations in ice-pellet conditions. 

The environmental chamber testing did not indicate that the imbedded ice pellets were conducive to 

contamination adhering to the underlying surface. Researchers decided to examine whether this was also 

true during actual takeoff runs, and also whether the presence of the ice pellets somehow interfered with 

the flowoff of the fluid during a takeoff run. Adhering contamination or a significant amount of fluid 

remaining on the wing at takeoff would both represent a safety hazard. 

The FAA Ground Icing Program and Transport Canada sponsored research conducted at the 

MacDonald Cartier International Airport, Ottawa, Canada, in March 2006 using a Dassault Falcon 20 

aircraft belonging to and operated by NRC Canada. APS Aviation, Inc. was contracted to support this 

research. Ice pellets were simulated using crushed ice from blenders which was run through sieves to get 

a size distribution similar to natural pellets. Type IV fluids were applied to a designated portion of the 

Falcon 20 wing followed by the simulated ice pellets that were distributed in known quantities and rates

using modified and calibrated hand held spreaders. The test segment of the wing was on the inboard half 

and had a trailing-edge flap, but no leading-edge slat. The flaps were not deployed until the start of the 

simulated takeoff. The Falcon 20 was operated through a simulated takeoff run excluding climbout. For 

most of the trials, the maximum speed ranged from 119 to 125 kn (61.2 to 64.3 m/s) that was attained in 

23 to 27 s after brake release. The contaminated fluid flowoff behavior was visually observed and 

recorded, and no aerodynamic performance data were reported. These tests indicated that when the 

aircraft was protected with Type IV anti-icing fluid, the fluid contaminated with the simulated pellets 

would readily flow off prior to rotation for all but the most extreme conditions tested, and that no frozen 

contamination remained on the wing. Review of the research indicated that a 25 min allowance time after 

anti-icing in light ice pellets afforded a large safety margin. The FAA approved a 25 min allowance time 

for winter 2006-2007 after anti-icing in light ice pellets with undiluted Type IV fluid, giving air carriers 

adequate time to get off the ground at most airports. An allowance time differs from a HOT in that there 

is no provision for pre-takeoff contamination check if the time is exceeded, meaning that the allowance 

time cannot be extended by a visual cabin check or external tactile or visual check of the aircraft critical 

surfaces within 5 min of takeoff. Ordinarily pre-takeoff contamination checks are conducted from the 

cabin, and the research showed that such a check would be of little use in the case of ice pellets (Ref. 4).

Ice pellets often occur mixed with other forms of precipitation in North America (Ref. 5). Thus the 

FAA recognized a need for further testing for various mixed conditions and temperature ranges. The 

following winter (January-February 2007), again in cooperation with Transport Canada, ice-pellet 

contamination testing was conducted in a large, open-circuit wind tunnel. This facility, the Propulsion and 

Icing Wind Tunnel (PIWT) is owned and operated by NRC. A two-dimensional, single element, NACA 

23012 airfoil model with 4.0 ft (1.2 m) chord was used for this test campaign (Ref. 6). An airplane takeoff 

acceleration profile and rotation was simulated in the wind-tunnel testing typically reaching a maximum 

speed of 100 kn (51.4 m/s) in 30 to 35 s. Allowance times were determined using the visual observations 

of the researchers. Limited aerodynamic measurements were also made, supplementing the observations 

of the researchers. However, the results did not show large effects of the contaminated fluids on the 

coefficient of lift, and no aerodynamic criteria were introduced for the determination of allowance times. 

The PIWT tests were also supplemented by environmental chamber tests whereby plates contaminated 

with ice pellets were be subjected to simulated freezing rain, freezing drizzle, and snow. Based on the 

results from the testing, the FAA published an allowance time table for winter 2007-2008 in Notice N 

8900.19 (Ref. 7) and also in Notice N 8900.22 (Ref. 8). The table is reproduced as Table I of this report. 

In February and March 2008, simulated takeoff testing (excluding climbout) was conducted using 

NRC’s Dassault Falcon 20 and Lockheed-Martin T-33 airplanes. The primary purpose of this testing was 

further evaluation of the allowance times already provided (cf. Table I) to determine if any changes were 

needed. Therefore, methods were developed to provide simulated ice-pellet conditions and mixed 

conditions (ice pellets mixed with snow, freezing rain, freezing drizzle) on the runway. The methods and 

procedures used in the March 2006 Falcon 20 tests were generally followed in this campaign. The

contaminated fluid flowoff behavior was visually observed and recorded, and no aerodynamic 
NASA/TM—2012-216014 4 

performance data were reported. In addition to the Falcon 20 inboard wing segment used in March 2006, 

fluid and contamination was also applied to an outboard section of the Falcon 20 wing. This allowed for 

observation of the effect of extended leading-edge slat on contaminated fluid flowoff behavior. The 

rotation speed was nominally 120 kn (61.7 m/s) at 25 s after brake release. The effect of rotation speed 

was also considered in this test campaign with trials conducted at a nominal speed of 80 kn (41.1 m/s) at 

17 s after brake release. There were no changes to the allowance time table for winter 2008-2009 as a 

result of this testing.

TABLE I.—ICE PELLET ALLOWANCE TIMES FOR WINTER 2007-2008 (REFS. 7 AND 8) 

OAT –5 °C or warmer,

min

OAT colder than –5 °C,

min

Light ice pellets 50 30 

Moderate ice pellets 25 10 

Light ice pellets mixed with light or moderate snow 25 Operations not authorized

Light ice pellets mixed with light or moderate freezing drizzle, or light 

freezing rain (operations not authorized below –10 °C OAT)

25 10 

(operations not authorized 

below –10 °C OAT)

Light ice pellets mixed with light rain (operations not authorized 

below 0 °C OAT)

25 

(operations not authorized 

below 0 °C OAT)

Operations not authorized 

below 0 °C OAT)

The research returned to the NRC PIWT again in January-March 2009 using a 6.0 ft (1.8 m) chord 

LS(1)-0417 airfoil model. This model had a span of 7.9 ft (2.4 m) and was mounted in the tunnel between 

two end plates to minimize three–dimensional effects. This model had a single-element slotted flap 

deployed at 15°. Allowance times were again determined primarily using the visual observations of the 

researchers. Aerodynamic measurements were made, supplementing the observations of the researchers.

However, the results did not show large effects of the contaminated fluids (never more than 5 percent loss 

of lift), and no aerodynamic criteria were explicitly introduced for the determination of allowance times. 

Based on the results, the FAA published an updated ice pellet allowance time table for Winter 2009-10 in 

Notice N 8900.98 (Ref. 9) reproduced here as Table II. An allowance time for light ice pellets mixed with 

rain was added, and changes were made to the rows for light ice pellets mixed with light snow and light 

ice pellets mixed with moderate snow. 

TABLE II.—ICE PELLET ALLOWANCE TIMES FOR WINTER 2009-2010 (REF. 9) 

OAT –5 °C 

and above,

min

OAT less than 

–5 to –10 °C,

min

OAT less than 

–10 °C,

min

Light ice pellets 50 30 30

Moderate ice pellets 25 10 10

Light ice pellets mixed with light or moderate 

freezing drizzle

25 10

Caution: No 

allowance times currently 

exist

Light ice pellets mixed with light or freezing rain 25 10

Light ice pellets mixed with light raina 25

Light ice pellets mixed with moderate raina 25

Light ice pellets mixed with light snow 25 15

Light ice pellets mixed with moderate snow 10 a

Limited to OAT of 0 °C and above

Tests were continued in the NRC PIWT in January-February 2010 using a thin, high-performance 

airfoil model rigged with a single-element slotted flap fixed at 20°. The thin, high-performance section was 

regarded as typical of wings on regional jet transport airplanes. This model had a chord of 6.0 ft (1.8 m) and 

a span of 7.9 ft (2.4 m) and was mounted in the tunnel between two end plates. Clark and MacMaster

(Ref. 10) provide further information regarding the January-February 2010 wind-tunnel testing conducted at 

the NRC PIWT. Aerodynamic measurements were again made, and effects on lift were significantly larger 

than in past years. Along with visual evaluation by the researchers, a criterion for percentage loss in lift at 8°

angle of attack was explicitly used. This raised several concerns that are discussed in detail in Section 3.0 of 

this report. Based on the results, the FAA published an updated ice pellet allowance time table for winter 
NASA/TM—2012-216014 5 

2010-2011 and 2011-2012, since it was not changed, on the FAA AFS-200 Ground Deicing Web site

(Ref. 11) reproduced here as Table III. The allowance times for propylene glycol (PG) based fluids were 

reduced or eliminated in some cases. It was noted that ethylene glycol (EG) fluids tend to flow off more 

readily than PG fluids, especially when contaminated and for temperatures lower than –10 °C. The 

allowance times were eliminated for light ice pellets and moderate ice pellets for OAT less than –10 °C for 

PG fluids for aircraft with rotation speeds less than 115 kn (59.1 m/s), and were reduced from 25 to 15 min

for moderate ice pellets for OAT greater than 0 °C for PG fluids.

TABLE III.—ICE PELLET ALLOWANCE TIMES FOR WINTER 2010-2011 (REF. 11) 

OAT –5 °C

and above, 

min

OAT less than

–5 to –10 °C,

min

OAT less than

–10 °C,

min

Light ice pellets 50 30 30a

Moderate ice pellets 25b 10 10a

Light ice pellets mixed with light or moderate freezing drizzle 25 10

Caution: No allowance 

times currently exist

Light ice pellets mixed with light freezing rain 25 10

Light ice pellets mixed with light rainc 25

Light ice pellets mixed with moderate rainc 25

Light ice pellets mixed with light snow 25 15

Light ice pellets mixed with moderate snow 10 a

No allowance times exist for PG fluids, when used on aircraft with rotation speeds less than 115 kn. (For these aircraft, if the fluid type is not 

known, assume zero allowance time.)

b

Allowance time is 15 min for PG fluids, or when the fluid type is unknown.

c

Operations NOT authorized below 0 °C

Tests were continued in the NRC PIWT in January-February 2011 using the same thin, highperformance wing model used the previous year. There were no changes to the allowance times tables for 

winter 2011-2012 based on the testing. The evolution of the allowance time tables since their introduction 

is depicted in Table IV, which employs a common format to facilitate comparison of all the tables.

TABLE IV.—EVOLUTION OF ALLOWANCE TIME TABLES (ALL TIMES ARE IN MINUTES)

Winter 2007-2008 and 2008-2009 Winter 2009-2010 Winter 2010-2011 and 2011-2012

OAT 

≥ 0 °C

0 °C ˃ 

OAT

≥ –5 °C

–5 °C ˃ 

OAT

≥ –10 °C

–10 °C

˃ OAT

OAT 

≥ 0 °C

0 °C ˃ 

OAT

≥ –5 °C

–5 °C ˃ 

OAT

≥ –10 °C

–10 °C

˃ OAT

OAT 

≥ 0 °C

0 °C ˃ 

OAT

≥ –5 °C

–5 °C ˃ 

OAT

≥ –10 °C

–10 °C 

˃ OAT

Light ice pellets 50 50 30 30 50 50 30 30 50 50 30 30b

Moderate ice pellets 25 25 10 10 25 25 10 10 25a 25 10 10b

Light ice pellets mixed 

with light or moderate 

freezing drizzle

25 25 10 0 25 25 10 0 25 25 10 0 

Light ice pellets mixed 

with light freezing rain 25 25 10 0 25 25 10 0 25 25 10 0 

Light ice pellets mixed 

with light rain 25 0 0 0 25 0 0 0 25 0 0 0 

Light ice pellets mixed 

with moderate rain 0 0 0 0 25 0 0 0 25 0 0 0 

Light ice pellets mixed 

with light snow 25 25 0 0 25 25 15 0 25 25 15 0 

Light ice pellets mixed 

with moderate snow 25 25 0 0 10 10 0 0 10 10 0 0 

a

Allowance time is 15 min for PG fluid.

b

Allowance time is 0 min for PG fluid with rotation speed less than 115 kn.

Shading indicates that more time is allotted than in previous table.

Shading indicates that less time is allotted than in previous table.
NASA/TM—2012-216014 6 

The dissemination of the wind-tunnel testing results and subsequent changes to the ice-pellet 

allowance time tables have led to concerns regarding the fidelity and applicability of the measured lift 

degradation in the ice-pellet contamination fluid testing at the NRC PIWT. Given the potential 

operational safety implications associated with the measured performance degradations, it is even more 

important that these concerns be addressed appropriately. The list of potential issues includes the 

following: 

• A two-dimensional model with endplates was used instead of a three-dimensional geometry more 

representative of an airplane wing. 

• Lift degradations were measured at αw = 8° on the two-dimensional model instead of at stall.

• Airplanes undergoing take-off accelerations and rotation experience ground effects that were not 

simulated in the PIWT testing. 

• The chord Reynolds number was small relative to most airplane wing sections. 

• The 6-ft (1.8 m) chord length model is small relative to most airplane wing sections.

• The large model size and proximity of the wind-tunnel walls resulted in uncorrectable effects on 

the aerodynamic data and performance.

• The rotation speed used in the PIWT tests was low relative to typical airplane rotation speeds.

The essence of these concerns is understanding how the performance degradations measured on the model 

in the NRC PIWT testing relate to an actual three-dimensional, full-scale airplane configuration. 

Therefore, some means of determining the significance of the measured lift degradations relative to a fullscale airplane is necessary.

It has been suggested that the measured lift degradation due to the uncontaminated, or pristine, fluids 

in the NRC PIWT tests could be compared to the loss in maximum lift that is associated with the 

Aerodynamic Acceptance Test (AAT). The AAT is described in SAE Aerospace Standard AS5900 

(Ref. 12) and was developed as a means to qualify deicing and anti-icing fluids. The AAT fluid

acceptance criterion for utilizing the flat-plate, boundary-layer displacement thickness, δ*, is based upon 

a correlation with loss in maximum lift on a Boeing 737-200ADV configuration derived from flight and 

wind-tunnel testing. Therefore, δ* data resulting from the AAT for a specific fluid correlate to a loss in 

maximum lift on the B737-200ADV airplane. Relating this information to the measured lift degradation 

in the corresponding NRC PIWT testing using the same fluid provides a means to “scale” the measured 

lift degradation in the NRC PIWT tests. In order to follow this approach, a detailed review of the research 

basis for the AAT must be conducted. It is important to understand how the AAT qualification criterion 

was established in order to understand how it may be applied to the present aerodynamic testing of antiicing fluids. 

Therefore, the purpose of this report is to review the research basis of the AAT and determine how it

may be applied to the present aerodynamic testing of anti-icing fluids in the PIWT. A detailed review of the

research basis of the AAT is performed in Section 2.0 of this report. Section 3.0 then provides a review of 

the anti-icing fluid aerodynamic testing conducted at the NRC PIWT with the uncontaminated fluid. These 

results were used in conjunction with results from the AAT performed for the same fluids to relate the 

aerodynamic degradation on the two-dimensional PIWT model to a B737-200ADV airplane. This 

relationship was then used to develop a lift loss criterion that could be used to establish allowance times for 

ice-pellet contamination. This report describes how this “scaling” method may address the concerns raised 

in regard to the NRC PIWT tests and implications for establishing ice-pellet allowance times.

2.0 Review of the Research Basis of the Aerodynamic Acceptance Test

2.1 Introduction

A major research program was conducted in the mid to late 1980s and early 1990s to characterize 

deicing/anti-icing fluid behavior and aerodynamic effects on transport airplanes. This work was led by 
NASA/TM—2012-216014 7 

Eugene Hill and Thomas Zierten of the Boeing Commercial Airplane Group and is described in 

References 13 to 16. This research led to the development of the AAT that is described in Ref. 12. A 

flight test was conducted to establish the fluid effects on a full-scale Boeing 737-200ADV airplane.

Following the flight test, wind-tunnel testing was conducted with both three-dimensional and twodimensional subscale models based upon the B737-200ADV in the NASA Glenn (then Lewis) Icing 

Research Tunnel (IRT). The icing tunnel test results were correlated to the flight trials to determine the 

applicability of subscale wind-tunnel testing for deicing/anti-icing fluid effects. The IRT testing was 

extended to include boundary-layer measurements on the two-dimensional model. These measurements 

were subsequently correlated to boundary-layer measurements on a flat plate model to form the basic 

principles of the AAT. This review will describe each of these research efforts in more detail. 

2.2 Flight Test Campaign 

The flight test campaign was conducted during January 11 to 20, 1988, at Kuopio, Finland. The 

experiments were led by Boeing and supported by the Association of European Airlines (AEA) and three 

fluid manufacturers. “The objectives of the flight test were to evaluate the aerodynamic effects of ground 

deicing/anti-icing fluids on a large jet transport and to establish a full-scale data base for extrapolating 

data from a planned, more comprehensive wind-tunnel investigation. The primary goal of the flight test 

was to obtain data for assessing the fluid effects on lift.” (Ref. 13) To carry out these objectives, a series 

of takeoffs were performed using an instrumented Boeing 737-200ADV airplane. In addition to the 

airplane and air data acquisition, video and photographic records of the fluid flow characteristics were 

also obtained. Only Type I and Type II fluids were tested. The fluids were applied using a two-step, inservice procedure: the wing and horizontal stabilizer were first deiced using a 50:50 mixture of Type 1 

fluid and hot water and then anti-iced with undiluted, cold Type II fluid. When testing was conducted 

with only Type I fluid, the fluid was applied cold and undiluted without the deicing step. A picture of 

fluid application is shown in Figure 1. Four fluids were tested: 

• Fluid 1: Hoescht Type I

• Fluid 2: Obsolete Type II (used because existing wind-tunnel data were available)

• Fluid 3: Hoescht Type II 

• Fluid 4: Union Carbide Canada Type II 

Figure 1.—Application of anti-icing fluid to B737-200ADV 

during flight test trials at Kuopio, Finland, in 1988, from Hill 

and Zierten (Ref. 13). Reprinted with permission of the 

American Institute of Aeronautics and Astronautics.
NASA/TM—2012-216014 8 

Figure 2.—Summary of fluid effects on lift coefficient from B737-200ADV 

flight test trials, from Hill and Zierten (Ref. 13). 

Overall, the flight test results showed that the lift is lower with the fluids applied to the wing relative 

to the clean, dry baseline. A summary of the fluid effects on lift coefficient for a liftoff attitude of 12° is 

shown in Figure 2. Hill and Zierten (Ref. 13) note that the liftoff attitude, αB = 12° was typical of the oneengine inoperative takeoff climb attitude for the B737-200ADV. They also estimated a ±1.5 percent 

uncertainty in the experimental data. The baseline lift coefficient for αB = 12° corresponded to 75 percent 

of CL,max for the clean, dry airplane. Flight test data were not gathered near CL,max because this would have 

required stalling the airplane close to the ground (Ref. 16). The data in Figure 2 illustrates the effects of 

the four fluids at various temperatures and for both the flaps 5 and flaps 15 configurations. In general, the 

lift penalties due to the fluid were higher at flaps 15 than flaps 5, thus illustrating the effect of airplane 

configuration. The lift loss for the Type I fluid (Fluid 1) at –10 °C was less than that of a Type II fluid 

(Fluid 3) also at –10 °C. Hill and Zierten (Ref. 13) reported, “Limited assessments of flow off time, liftoff 

speed and fluid exposure time revealed that reasonable variations of these parameters had no measurable 

effects on the test results.” Visual observations of the fluid flowoff characteristics were also performed 

and led to various hypotheses about the behavior of the instability waves in the fluids. Detailed discussion 

of fluid flowoff characteristics and instabilities is beyond the scope of this report.

Hill and Zierten (Ref. 13) reported the following general conclusions from the B737-200ADV flight 

test trials:

1. “The neat fluids cause a measurable loss in lift, similar to the results of earlier Boeing and AEA 

wind-tunnel tests.

2. The fluid effects are configuration dependent. 

3. Effects of the fluids on airplane handling qualities can be small. 

4. A secondary fluid wave flows aft immediately after rotation for takeoff.

5. Fluid effects are transitory.”

References 13 to 16 provide more information supporting these conclusions. 
NASA/TM—2012-216014 9 

2.3 Phase I IRT Test Campaign

Both Phase I and II IRT tests were cooperative efforts among Boeing who provided the models and 

conducted the test, NASA Glenn who provided use of the IRT and test support, various fluid 

manufacturers who provided the deicing/anti-icing fluids and the AEA who monitored the test to maintain 

continuity with their fluids research program at that time. The first IRT test campaign, Phase I, was 

conducted in early 1988, shortly after the conclusion of the flight test trials. According to Hill and Zierten, 

(Ref. 13) the primary objective of the Phase I test campaign was to validate the use of wind-tunnel tests 

for investigating the aerodynamic effects of deicing/anti-icing fluids. 

The Phase I IRT campaign was conducted using an existing B737-200ADV model. This model was a 

9.1 percent scale, half-plane model mounted to a splitter wall along the tunnel test-section sidewall as 

shown in Figure 3. The forces and moments on the model were measured with a balance located in the 

splitter wall. The average chord length for this model was 1.0 ft (0.305 m) and the wing was equipped 

with a leading-edge slat and a trailing-edge flap. The model was tested with and without a ground plane 

installed in the test section as shown in Figure 3. The detailed test procedures are described in References

13 and 16, and are summarized here. After the fluid application, a 5-min exposure time was used to allow 

the fluid temperature to equilibrate with the model and air temperature. The wind-tunnel flow velocity 

was linearly increased to 135 KEAS (69 m/s) in approximately 30 s. After 25 s of tunnel flow, the model 

was rotated from the taxi attitude (αB = 0°) to the desired angle of attack at rotation rate of 3°/s. The speed 

at rotation was approximately 120 to 125 KEAS (62 to 64 m/s), corresponding to typical rotation speeds 

for the full-scale B737-200ADV airplane. A comparison of the flight test and IRT test acceleration profile 

is shown in Figure 4. The desired attitude varied, but was up to and including wing stall in many cases.

Using this method, the entire lift curve could be determined with a single run. This method for obtaining 

the entire lift curve in a single run was validated by performing a series of runs where the model was 

rotated to a fixed attitude to establish the lift curve in fixed increments. The aerodynamic performance 

data from both methods were comparable (Ref. 16).

Through the course of the initial IRT investigations, Runyan et al. (Ref. 16) considered the effect of 

the rotation speed, time to rotation and initial fluid depth on the aerodynamic performance. The effect of 

independent variations of these parameters on lift coefficient is shown in Figure 5 for Fluid 3 at αB = 7°. 

The first plot shows that there is an effect of the velocity at rotation with higher rotation speeds resulting 

in higher lift coefficients. The second plot shows that there was virtually no difference in lift coefficient 

for two different rotation times. In this case the rotation velocity was constant and the times were changed 

by changing the acceleration profile. Finally, the effect of the initial fluid thickness on lift coefficient was 

inconclusive due to the scatter in the data. Analogous plots for the percent increase in drag coefficient are 

shown in Figure 6. The drag data tend to follow the trends in the lift effects, with the possible exception 

of the effect of rotation time. The longer rotation time seemed to reduce the percent increase in drag 

coefficient slightly, but had little or no effect on lift coefficient. Based upon these results and the lift 

comparisons to be described in the following paragraph, no adjustment of the acceleration conditions was 

found to be necessary. That is, the time to rotation, rotation speed and rotation rate used in the IRT tests 

matched that used in the flight test campaign. 

Runyan et al. (Ref. 16) provide a good discussion of the limitations involved in the fluids-related 

wind-tunnel testing. They discussed the tradeoff and effects of model scale, Reynolds number, local flow 

velocities and surface shear stresses, etc. It was certainly acknowledged that the small model scale 

(9.1 percent) and associated Reynolds number effects would result in serious questions about significance 

and applicability of the wind-tunnel data. That is why Phase I was conducted to gain confidence that the 

wind tunnel could be used to evaluate fluid effects on aerodynamic performance. Hill and Zierten wrote 

(Ref. 13): “By testing the same deicing/anti-icing fluids used during the flight test program on the B737-

200ADV three-dimensional model, the correlation between full- and model-scale fluid effects on lift...was 

obtained.” The authors refer to Figure 7 that shows the comparison between percent lift loss on the 
NASA/TM—2012-216014 10

Figure 3.—9.1 percent scale, three-dimensional half-plane B737-200ADV model installed in NASA IRT, from 

Runyan et al. (Ref. 16). 

Figure 4.—Comparison of takeoff acceleration profiles between the flight 

test trials and the IRT test campaign, from Runyan et al. (Ref. 16). 
NASA/TM—2012-216014 11

Figure 5.—Effect of acceleration parameters on lift coefficient at αB = 7° on the 9.1 percent scale, three-dimensional

half-plane B737-200ADV model with sealed slat, flaps 5 configuration, Fluid 3 at –10 °C air temperature with 

ground plane installed, from Runyan et al. (Ref. 16). 

Figure 6.—Effect of acceleration parameters on percent increase in drag coefficient at αB = 7° on the 9.1 percent

scale, three-dimensional half-plane B737-200ADV model with sealed slat, flaps 5 configuration, Fluid 3 at -10 °C air 

temperature with ground plane installed, from Runyan et al. (Ref. 16). 

Figure 7.—Comparison of fluid effects on percent lift loss 

between the full-scale B737-200ADV airplane and the 9.1

percent scale, three-dimensional half-plane model in the IRT 

tests, from Hill and Zierten (Ref. 13). 
NASA/TM—2012-216014 12

full-scale airplane from the flight test trials and the 9.1 percent scale, three-dimensional half model in the 

IRT. An important note to Figure 7 is that the lift loss on the scale IRT model was calculated at αB = 7°

instead of the αB = 12° used for the flight test data. The choice of αB = 7° on the subscale model was made 

because this corresponded to 75 percent of the CL,max on the clean model. The choice of 75 percent CL,max

as the comparison point was selected since the takeoff safety speed condition (one-engine-out climb) 

corresponds to about 75 percent of CL,max. Presumably, αB = 7° on the subscale model was lower than αB = 

12° on the full-scale airplane due to Reynolds number effects. Hill and Zierten then came to this 

conclusion: “Realizing the estimated ±1.5 percent data accuracy, the correlation [Figure 7] suggested 

that the wind-tunnel lift loss results could be extrapolated to full scale on a 1:1 basis.” The authors also 

go on to point out that similar results have been observed with distributed solid roughness on the wing. 

Therefore, the authors showed that the wind-tunnel testing of deicing/anti-icing fluids on a 9.1 percent

scale, three-dimensional half-plane B737-200ADV model produced results that were comparable to the 

full-scale flight test without applying any scale factors or other corrections. This allowed for the expanded 

matrix of fluids effects testing that was conducted in the phase II IRT test campaign. 

2.4 Phase II IRT Test Campaign

Having addressed the validity of the IRT testing with the three-dimensional half-plane model in Phase 

I, the test objectives were expanded in Phase II to include:

• Determining the effects of the fluid on maximum lift.

• Expanding the testing envelope in terms of fluids used, temperature ranges, aerodynamic and 

fluid-flow measurements.

• Acquisition of boundary-layer displacement thickness data that were subsequently used in the 

development of the AAT. 

In order to accomplish these objectives further tests were conducted with the 9.1 percent scale, threedimensional half-plane model described in previous section and illustrated in Figure 3. In addition, a twodimensional model test set up was also introduced into the IRT test section. The two-dimensional model 

was an 18 percent scale (1.5 ft (0.457 m) chord) section of the B737-200ADV wing at 65 percent

semispan station. It was mounted horizontally between two splitter walls as shown in Figure 8 and no 

ground plane was used. This section (at 65 percent semispan) was selected because it was “critical” 

relative to the stall progression on the baseline, clean, full-scale B737-200ADV airplane. The 

aerodynamic performance data were acquired from a force and moment balance mounted in the splitter 

walls. In addition to these measurements, the boundary-layer velocity profile was measured with a 

pressure rake mounted toward the aft end of the model, just upstream of the flap. The velocity profile was 

integrated to obtain the boundary-layer displacement thickness.

Since the flight test trials were conducted in close proximity to the ground, safety considerations 

prevented measurements of fluid effects near stall. That is why the lift loss results were reported at αB = 

12°. Determining the fluid effects on maximum lift was accomplished safely in the wind tunnel and the 

results for a variety of fluids and air temperatures are shown in Figure 9 for the 9.1 percent scale, threedimensional half-plane B737-200ADV model. These data show that percent lift loss was much larger at 

CL,max that at αw = 8°. (Note that αw = 8° corresponds exactly to αB = 7°) These large differences 

emphasized the importance of evaluating the effect of the fluids at stall. Since no full-scale data were 

available for the effects of the fluids on CL,max, it had to be assumed that the subscale model IRT results at

CL,max could be extrapolated to full-scale on a one-to-one basis. In further support of this assumption, 

Runyan et al. (Ref. 16) present aerodynamic data acquired on the 9.1 percent scale, three-dimensional 

half-plane B737-200ADV model with distributed solid roughness and simulated frost in the IRT. It was 

noted that the percent reductions in maximum lift agreed closely with flight test data for similar roughness 

configurations. Runyan et al. (Ref. 16) also describe a number of parametric studies designed to ascertain 
NASA/TM—2012-216014 13

Figure 8.—18 percent scale, two-dimensional model of the B737-200ADV wing section installed in the NASA IRT, 

from Runyan, et al. (Ref. 16). 
NASA/TM—2012-216014 14

Figure 9.—Summary of the deicing/anti-icing fluid effects on percent lift loss at αw = 8° (αB = 7°) and at 

CL,max for the 9.1 percent scale, three-dimensional half-plane model in the NASA IRT, from Hill and 

Zierten (Ref. 13). 

the effect of various slat and flap configurations on fluid flowoff and resulting aerodynamic performance 

effects. While it is beyond the scope of this report to review these results in detail, they were carefully 

analyzed to support the hypotheses regarding the fluid flow behavior and aerodynamic degradation. 

Having concluded that the 9.1 percent scale, three-dimensional half-plane model results were 

representative of full-scale on a one-to-one basis, the next step was to evaluate lift degradation on the 18

percent scale, two-dimensional wing section. Numerous parametric studies were performed looking at the 

effects of the various fluids, air temperature and other factors. The lift degradation was measured at both 

αw = 8° (αB = 7°) and at stall. Some sample results are shown in Figure 10. Note here that Cl,max (with 

lower case “l”) is the two-dimensional sectional maximum lift coefficient. Hill and Zierten (Ref. 13) 

commented, “Results shown in [Figure 10] for the two-dimensional model, indicate percentage lift losses 

at 8° angle of attack that are larger than those at maximum lift. This trend is different from that observed 

on the three-dimensional model and indicates the importance of three-dimensional testing.” Presumably, 

the authors meant that the importance of three-dimensional testing is relative to determining accurately 

the effects of the fluids at maximum lift. This is an important point because the authors conducted 

numerous other tests with the two-dimensional model. Runyan et al. (Ref. 16) offer a clarification on this 

point: “...these two-dimensional results are useful for determining the relative fluid-to-fluid lift losses at 

a given temperature and temperature-to-temperature lift losses for a given fluid. However, since these are 

two-dimensional data, they cannot be used directly to estimate lift losses on the airplane.” Therefore, the 

two-dimensional model testing was useful for evaluating the relative effects of the fluids, temperature, 

contamination and other factors. This conclusion suggests that meaningful relative comparisons of fluid 

aerodynamic effects can be obtained in two-dimensional model testing. However, Hill and Zierten

(Ref. 13) suggested that accurate absolute percentage lift losses were only obtained using the threedimensional half-plane model. 
NASA/TM—2012-216014 15

Figure 10.—Effect of fluid temperature, dilution and contamination (simulated precipitation) on lift loss at

αw = 8° and at Cl,max for the 18 percent scale two-dimensional model in the NASA IRT, from Hill and 

Zierten (Ref. 13). 

Through the course of this and previous investigations, hypotheses relating the fluid-flow-off 

behavior to the aerodynamic degradation were being developed and tested. It is beyond the scope of this 

review to describe these hypotheses, related experiments and observations with the exception of the 

hypothesis leading to the development of the AAT. Hill and Zierten (Ref. 13) suggested, “If the 

hypothesis is valid that the mechanism of the fluid effect on the aerodynamics of the wing is premature 

growth of the boundary layer, then a correlation between the prematurely thickened boundary layer and 

the aerodynamic effect can be anticipated.” The authors quantified the boundary-layer displacement 

thickness through direct measurement and subsequent integration of the boundary-layer velocity profile.

Their hypothesis was then supported by Figure 11 that illustrates the suggested correlation between the 

percent lift loss on the two-dimensional model at αw = 8° and δ* measured on the same model at the same 

condition. Since the data for the various fluids and air temperatures “collapse” onto the correlation curve, 

the conclusion is that the hypothesis is valid. Hill and Zierten continue: “Because the two-dimensional 

model airfoil is the critical (relative to stall progression) 737-200ADV wing section and assuming that 

the thickened boundary-layer measured at 8-deg. angle of attack was an indicator of the effect of fluid at 

maximum lift, then an acceptable correlation between the two-dimensional model boundary-layer 

thickness and the three-dimensional model maximum lift loss can also be anticipated.” The suggested 

correlation is shown in Figure 12 that plots the three-dimensional model percent loss in maximum lift 

against δ* measured on the two-dimensional airfoil model at αw = 8° for the same fluids at the same 

conditions used on each model. All of the data for the various fluids at three different temperatures tend to 

collapse onto the curve shown in Figure 12, suggesting that the correlation is, as suggested, acceptable.

Therefore, Hill and Zierten concluded that these results indicate that the fluid effects on the boundarylayer thickness (on the two-dimensional model at αw = 8°) is an indicator of fluid effects on airplane 

performance.
NASA/TM—2012-216014 16

Figure 11.—Correlation between percent lift loss and boundary-layer displacement 

thickness as measured on the two-dimensional model at αw = 8° for various fluids at air 

temperatures of 0, –10 and –20 °C, from Hill and Zierten (Ref. 13). 

Figure 12.—Correlation between three-dimensional B737-200ADV model percent 

maximum lift loss and two-dimensional model boundary-layer displacement thickness 

at αw = 8° for various fluids at air temperatures of 0, –10 and –20 °C, from Hill and 

Zierten (Ref. 13). 
NASA/TM—2012-216014 17

These correlations and the data plotted in Figure 11 and Figure 12 imply a direct relationship between 

the percent maximum lift loss measured on the three-dimensional model and the percent lift loss 

measured at αw = 8° on the two-dimensional model for the same fluids tested on each model under 

identical conditions. These data were extracted from Figure 11 and Figure 12 and are plotted in Figure 13. 

This could be done because the δ* data plotted on the horizontal axis was common to both plots. In 

Figure 11, a measured value of δ* for a given fluid at a given temperature was associated with a percent 

lift loss on the two-dimensional model at αw = 8°. The same fluid at the same temperature was tested on 

the three-dimensional model as well, providing a specific percent loss in maximum lift. Thus, the same δ* 

measured in the two-dimensional test for the same fluid at the same temperature was correlated with

percent maximum lift loss on the three-dimensional model, even though δ* was not measured on the 

three-dimensional model. The resulting correlation is shown in Figure 13 and a linear fit was constructed 

through the data. The y-intercept was forced through the origin because this is the limiting case for the dry 

wing. That is, the dry wing has no lift loss for either the two-dimensional or three-dimensional model, by 

definition. The slope of the line, 1.04 is very close to one which is consistent with the statement by Hill 

and Zierten that the two-dimensional wing section tested was the critical section relative to the stall 

progression. The data show that the implied correlation between the lift loss on the two-dimensional 

model at αw = 8° and the maximum lift loss on the three-dimensional model was better for some fluids 

than for others. Note that there are three data points for each fluid corresponding to air temperatures of 0, 

–10 and –20 °C as in Figure 11 and Figure 12. Generally, the percent lift loss increases with decreasing 

temperature. The data for the Dow, UCAR AAF and SPCA fluids tend to match the correlation fairly 

well, whereas the data for the Hoeschst and Kilfrost fluids clearly had a different effect between the twodimensional and three-dimensional models. This same “scatter” is represented in Figure 11 and Figure 12, 

however, the discrepancies do not appear to be as large due to the choice of scales on the plots. These data 

illustrate the potential variation in fluid behavior and their attendant effects on aerodynamic performance.

Figure 13.—Correlation between percent lift loss on the two-dimensional model at αw = 8° and percent maximum lift 

loss on the three-dimensional B737-200ADV model for a given fluid at a given temperature in the Phase II IRT 

tests. Note that there are three data points for each fluid representing 0, –10 and –20 °C. Data extracted from Hill 

(Ref. 14) 

⊗

⊗

⊗

2D Model Lift Loss at α = 8 deg., %

3D B737-200ADV Model Maximum Lift Loss, %

0.0 1.0 2.0 3.0 4.0 5.0 6.0 7.0 8.0 9.0 10.0

0.0

1.0

2.0

3.0

4.0

5.0

6.0

7.0

8.0

9.0

10.0

11.0

Dry

Hoeschst

Dow

UCAR AAF

UCAR ADF

SPCA

Kilfrost

Octagon

Linear Fit

⊗

y = 1.04x
NASA/TM—2012-216014 18

Hill and Zierten (Ref. 13) summarized the main conclusions from both the Phase I and Phase II IRT 

test campaigns:

1. “Fluid effects can be evaluated in the wind tunnel and the results can be extrapolated to full 

scale with reasonable confidence. 

2. Dependent on temperature, Newtonian Type I and non-Newtonian Type II fluids produce 

measurable adverse aerodynamic effects.

3. For the B737-200ADV three-dimensional configuration, lift loss at maximum lift are greater than 

that at lower angles of attack.

4. Fluid aerodynamic effects are transitory.

5. The fluid effects vary with airplane configuration. 

6. Aerodynamic effects of the fluids correlate well with boundary-layer thickening caused by the 

fluids.

7. Aerodynamic effects of the fluid vary with specific fluids and viscosity is not the only governing 

rheological parameter.”

2.5 Basis of the AAT

The research and findings regarding the potential transitory aerodynamic degradation due to fluids led 

an international working group to develop the AAT for the fluids. This working group was formed under 

the Aerospace Industries of America Transportation Committee and consisted of industry representatives 

and others members. This AIA TC 218-4 group worked with the von Karman Institute and the AEA to 

develop the AAT. Research performed at the von Karman Institute showed that the fluid flowoff behavior 

on a small flat plate in a wind tunnel capable of operational airspeeds and temperatures correlated to the 

flowoff characteristics observed on the two-dimensional B737-200ADV wing section tested in the NASA 

IRT. This correlation, plotted in Figure 14, suggests a quasi-linear relationship between the displacement 

thickness measured on the two-dimensional airfoil model and the flat plate testing that was conducted at 

the von Karman Institute. 

Figure 14.—Correlation between displacement thickness measured on the two-dimensional

B737-200ADV wing section in the IRT and the displacement thickness measured on the 

flat plate in the von Karman Institute experiment, from Hill and Zierten (Ref. 13).
NASA/TM—2012-216014 19

Figure 15.—Correlation between maximum lift loss on the full-scale B737-200ADV 

airplane and the displacement thickness measured on the flat plate at von Karman 

Institute, from Hill and Zierten (Ref. 13). 

The typical flat-plate test apparatus and wind-tunnel requirements currently used in the AAT is 

detailed in Reference 12. In this case, the deicing/anti-icing fluid is applied uniformly to the horizontal 

floor of a rectangular duct having a width of 1.0 ft (0.3 m) and sufficient height and length to satisfy the

dry-plate boundary-layer displacement thickness specifications. Operational temperatures of the plate, air 

and fluid are also maintained at the desired set point. The wind-tunnel airflow over the plate is accelerated 

to approximately 127 kn (65 m/s) in 25 s to simulate the takeoff profile. The displacement thickness is 

measured after 30 s at a location 5.0 ft (1.5 m) downstream from the leading edge. The displacement

thickness is measured indirectly using static pressures to determine the local flow velocity, relating this to 

an area change and subsequently the displacement thickness. This is different from the more direct 

method measuring the boundary-layer velocity profile used in the IRT experiments. However, the details 

of the measurement methods should not affect the results if performed properly. 

The series of correlations presented in Figure 11, Figure 12 and Figure 14 imply that there exists a 

final correlation between the maximum lift loss measured on the 9.1 percent scale three-dimensional 

B737-200ADV half-plane model in the IRT and the boundary-layer displacement thickness measured on 

the flat plate. This final correlation is shown as a linear relationship in Figure 15. To summarize, this 

correlation was developed through the following steps: 

• Takeoff profile flight tests were performed on a B737-200ADV airplane to measure the effects of 

the fluid on lift. This was specifically documented at αB = 12°, corresponding to 75 percent of 

CL,max on the clean, dry airplane. 

• Subsequent tests performed in the NASA IRT on a 9.1 percent scale, three-dimensional B737-

200ADV half-plane model using the same fluids and temperatures as in the flight tests showed 

that percentage lift losses at 75 percent CL,max (corresponding to αB = 7° on the subscale model) 

could be extrapolated to full-scale on a one-to-one basis (Figure 7).

• Furthermore, aerodynamic testing on subscale models and flight tests with simulated roughness 

and frost indicated that the percent maximum lift losses (i.e., ΔCL,max) measured on the 9.1 percent

scale three-dimensional B737-200ADV half-plane model could also be extrapolated to full-scale 

on a one-to-one basis. 
NASA/TM—2012-216014 20

• Therefore, the effect of fluids on full-scale airplane maximum lift was determined directly from 

IRT tests conducted on the 9.1 percent scale three-dimensional B737-200ADV half-plane model. 

• IRT tests performed with the same fluids and temperatures on the 18 percent scale twodimensional B737-200ADV wing section model showed that percent lift losses measured at 

αw = 8° correlated with:

○ Boundary-layer displacement thickness on the 18 percent scale two-dimensional B737-

200ADV wing section at αw = 8° (Figure 11) 

○ Percent maximum lift loss on the 9.1 percent scale three-dimensional B737-200ADV halfplane model (Figure 13).

○ Percent maximum lift loss on the 9.1 percent scale three-dimensional B737-200ADV halfplane model correlated with the boundary-layer displacement thickness on the 18 percent 

scale two-dimensional B737-200ADV wing at αw = 8° (Figure 12).

• The displacement thickness measured on a small-scale flat plate at the von Karman Institute with 

the same fluids and temperatures correlated with the displacement thickness measured on the 

18 percent scale two-dimensional B737-200ADV wing section at αw = 8° (Figure 14).

• Therefore, the percent maximum lift losses on the full-scale B737-200ADV airplane correlated 

with the displacement thickness measured on the flat plate (Figure 15.) 

This summarizes the development of the AAT for deicing/anti-icing fluids. The research showed that it

was possible to conduct meaningful small-scale, flat-plate experiments to evaluate the potential 

aerodynamic effects of these fluids. The remaining step in the development of the AAT was to establish 

the acceptance criterion.

2.6 Fluid Acceptance Criterion

Having established a relatively low-cost and simple means of evaluating the fluid effects on 

performance through the flat-plate boundary-layer displacement thickness, Hill and Zierten developed an 

acceptance criterion based upon the allowable maximum lift penalty on the B737-200ADV airplane. This 

is described in References 13 and 15 and involved an evaluation five specific takeoff performance criteria 

following FAR 25 requirements. Hill and Zierten determined that the most critical of these five criteria 

was maintaining: adequate margin between the 1g stall speed and the takeoff safety speed, V2. Hill and 

Zierten continue in Reference 15: “To insure safe operation when the fluids are used, a criteria requiring 

V2 speed to be 1.1Vs1g or greater was selected. The 10 percent V2 speed margin to the 1g stall speed 

compares with a 13 percent margin requirement interpreted from FAR 25.107(b)(1) for the nominal dry, 

clean wing. For aircraft whose V2 speed is at the minimum margin, 1.13 Vs1g, the acceptable lift loss 

resulting from the fluids is 5.24 percent.” The algebraic analysis used to arrive at the 5.24 percent 

reduction in CL,max is paraphrased from Reference 15: 

V2

Dry ≥ 1.13Vs1g (1) 

V2

Fluid ≥ 1.10Vs1g (2) 

∆CL,max

CL,max

Dry = CL,max

Dry −CL,max

Fluid

CL,max

Dry (3) 

∆CL,max

CL,max

Dry = 1 − CL,max

Fluid

CL,max

Dry (4) 
NASA/TM—2012-216014 21

∆CL,max

CL,max

Dry ≤ 1 − �1.10Vs1g�

2

�1.13Vs1g�

2 (5) 

∆CL,max

CL,max

Dry ≤ 5.24% (6) 

So, the 5.24 percent maximum lift loss (ΔCL,max) defines an acceptance criterion based upon the minimum 

margin for the V2 speed to 1.10Vs1g. It is important to note again that this is specific to the B737-200ADV 

used in this development. There is reduced concern for airplanes with larger margin for the V2 speed. 

Returning to the context of the AAT, the 5.24 percent maximum lift loss criterion is shown in 

Figure 15 as the upper limit in lift loss and corresponding boundary-layer displacement thickness 

measured in the flat-plate experiment at von Karman Institute. The plot shows that the 5.24 percent cutoff correlates δ* ≈ 9.2 mm for the fluid at –20 °C. Figure 15 also indicated that there are slightly lower 

maximum lift loss criteria for temperatures of –10 and 0 °C. Hill and Zierten (Ref. 13) explained that this 

was imposed because runways would be more likely to be contaminated at these warmer temperatures.

Presumably the contaminated runways would also have an adverse impact on the airplane performance in 

addition to the presence of the deicing/anti-icing fluid. While this may be an important consideration for 

the AAT, it is important to realize the that reduced acceptance criteria (either in terms of ΔCL,max or δ*) 

was not based upon a fluid-specific requirement, but concerns about additional adverse effects due to 

contaminated runways.

2.7 The AAT—Conclusions

This brief summary of the research basis of the AAT has shown how the potential adverse 

performance effects of deicing/anti-icing fluids on a B737-200ADV airplane were related to the 

boundary-layer displacement thickness measured on a flat plate. The relatively simple and low-cost flatplate experiment has subsequently been used as a part of the qualification process for ground deicing/antiicing fluids. As noted in the previous sections, the aerodynamic effects of these fluids are configuration 

dependent and can be influenced by other factors such as the takeoff speed. Quoting Hill and Zierten

(Ref. 15) again, “Compliance with the acceptance test is considered a minimum requirement since the 

test’s acceptance criterion is derived from fluid effects on a specific aircraft design [B737-200ADV] and 

only considers adequate takeoff safety speed margins. An airframe manufacturer may impose additional 

requirements which reflect considerations for specific airplane designs and performance criteria not 

addressed by the acceptance test.” 

3.0 Aerodynamic Evaluation of Anti-Icing Fluids in the NRC PIWT

3.1 Introduction

As described in Section 1.0, a cooperative effort among FAA, Transport Canada, NRC and APS has 

been undertaken to develop allowance times for ice-pellet conditions. The aerodynamic testing of the 

fluids with and without contamination was conducted at the PIWT facility on the NRC Montreal Road 

campus in Ottawa. One of the criteria used to evaluate the effectiveness of the fluid protection is the lift 

loss caused by the fluid/contamination as measured in the wind tunnel. The use of a lift loss performance 

criterion requires the establishment of some threshold value above which some safety level would be 

exceeded. (Such a threshold was developed for the AAT as described in Sec. 2.6.) Since the testing has 

employed a generic, thin, high-performance wing section, it is unclear what the appropriate lift loss 

threshold should be. Some means of relating the wind-tunnel model lift degradation to a real airplane 

configuration is needed in order to determine a reasonable lift loss threshold. This is important because of 

the potential operational safety implications associated with the measured performance degradations due 

to contaminated fluids. 
NASA/TM—2012-216014 22

It has been suggested that the measured lift loss due to the uncontaminated fluid in the PIWT tests can 

be compared to the B737-200ADV lift loss that is related to the AAT (as will be explained in this 

section). This approach provides a means to “scale” the measured lift degradation in the PIWT tests to 

equivalent maximum lift degradation on a full-scale B737-200ADV airplane. As reviewed in Section 2.5, 

Hill and Zierten (Ref. 13) showed that two-dimensional airfoil percent lift loss and boundary-layer 

displacement thickness correlated with the three-dimensional model maximum lift loss. Therefore, the 

thin, high-performance PIWT wing model lift loss may be correlated with the AAT results for tests with 

the same uncontaminated fluids at the same temperatures. Since the AAT results are directly related to 

percent maximum lift loss on the B737-200ADV airplane this correlation can be applied to establish an 

equivalent lift loss threshold for the thin, high-performance wing model in the NRC PIWT tests. 

The purpose of this section is to demonstrate how this analysis was carried out. Samples of the fluids 

used during the 2010 and 2011 PIWT test campaigns were subjected to the AAT as described in 

Reference 12 at the University of Quebec at Chicoutimi Anti-Icing Materials International Laboratory 

(AMIL). The AAT was performed over temperature ranges that closely matched those used in the PIWT 

test campaigns. For each fluid, the boundary-layer displacement thickness variation with temperature was 

quantified according to the Reference 12 procedures. An analysis was then performed to determine the 

relationship between the measured AAT boundary-layer displacement thickness, the PIWT model percent 

lift degradation for the fluid and the corresponding percent maximum lift loss on the B737-200ADV 

airplane. This analysis was used to define a lift loss threshold for evaluating the fluid performance on the 

model in the PIWT test. This procedure and results are described in the Sections 3.3, 3.4, and 3.5. The 

experimental methods used in the PIWT tests are described first in Section 3.2. 

3.2 PIWT Test Experimental Methods

The experimental methods used to evaluate the performance of fluids in the uncontaminated state and 

with ice-pellet contamination has been described elsewhere, so only a brief overview is given here. For 

example, Clark and MacMaster (Ref. 17) also provide a detailed description of the test set up, procedures

and results. The PIWT is an open-return wind tunnel that draws air from the outdoors and is therefore 

naturally cooled. The fan is located in the wind-tunnel inlet and was driven by a gas turbine engine for 

these test campaigns. The test section measures 10 ft (3 m) across and 20 ft (6 m) in height. For these tests 

an insert was used that reduced the height to 16.4 ft (5 m) resulting in an increase in maximum test 

section speed. A two-dimensional airfoil section was mounted horizontally in the test section as shown in 

Figure 16. The model had a span of 7.9 ft (2.4 m) and was isolated from the wind-tunnel walls with end 

plates. Due to the large gaps between the end plates and the test-section walls (≈ 1 ft (0.3 m) on each side) 

and the small size of the end plates, the model installation is quasi-two-dimensional. Figure 17 shows the 

6 ft (1.8 m) chord, thin, high-performance wing section with single-slotted flap deflected at 20°. The 

model had a fixed (or “hard”) leading edge and was regarded as typical of wings on regional jet transport 

airplanes. The model was supported by a force balance at each end located outside of the test section 

walls. A drive motor and gearbox connected to the force balance provided the rotation for setting angle 

of attack. The lift data and angle of attack were corrected for wind-tunnel wall effects using the 2D airfoil 

methods described in Rae and Pope (Ref. 18).

Data from a typical simulated takeoff run are shown in Figure 18. Plotted is the model lift coefficient 

versus time. These data were acquired by accelerating the test section airflow from 20 kn (10.3 m/s) to 

100 kn (51.4 m/s) in less than 31 s. The model was rotated from αw = –2° to 8° when the speed reached 

100 kn. The rotation rate was 2.7°/s. The wing was held at 8° for 1.5 s to gather more data at this angle 

of attack, then the angle was reduced. This was a typical profile. In other runs, the angle of attack was 

increased through stall along with other variations in rotation speed and time. The large rate of increase 

in Cl in Figure 18 corresponds to the model rotation from –2° to 8° angle of attack. During post 

processing, the Cl vs. αw data were extracted and plotted as shown in Figure 19. There are data for three
NASA/TM—2012-216014 23

Figure 16.—Photographs of thin, high-performance wing model installed in the NRC PIWT facility.

Figure 17.—Thin, high-performance airfoil section used in the PIWT testing, reference chord = 6.0 ft (1.8 m). 

Figure 18.—Sample lift coefficient vs. time profile for the simulated takeoff acceleration and wing rotation 

used in the PIWT tests.

x/c

y/c

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0 1.1

-0.10

-0.05

0.00

0.05

0.10

Time, sec.

PIWT Model Lift Coefficient,

Cl

10.0 15.0 20.0 25.0 30.0 35.0 40.0 45.0 50.0 55.0 60.0 65.0

0.0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

1.6

1.8

2.0

Lift coefficient data

during rotation from

αw = -2 deg. to 8 deg.

used for analysis.
NASA/TM—2012-216014 24

Figure 19.—Sample lift coefficient versus angle of attack data for three configurations—dry wing, uncontaminated

fluid only and fluid with ice-pellet contamination.

configurations: the dry wing, the wing with uncontaminated fluid only and the wing with fluid plus icepellet contamination. These data illustrate typical deicing/anti-icing fluid effects on lift. The lift curves for 

the fluid only and fluid with ice pellets are shifted below and to the right of the “Dry Wing” lift curve.

This effect is due to the presence of the fluid and contamination on the wing upper surface. The effect is 

to reduce the wing camber causing a shift in the zero-lift angle of attack. This effectively makes the 

profile more symmetric—and a symmetric airfoil has zero lift at zero angle of attack. The fact that the lift 

curves appear to be nearly parallel is not uncommon in fluids testing and has been reported in other 

studies (Ref. 16). It was determined that the percent lift loss on the model would be evaluated at αw = 8°. 

As shown in Figure 19, the fluid only configuration resulted in a 4.2 percent loss in lift from the dry wing 

at αw = 8 °. The addition of ice pellets to the fluid resulted in a further reduction in lift coefficient of 8.0 

percent from the dry wing configuration. The objective of the analysis described in this report is to 

determine the significance of these percent lift loss values relative to a full airplane configuration.

3.3 Summary of 2010 and 2011 Uncontaminated Fluid Tests at NRC PIWT

A series of simulated takeoff runs were performed in both 2010 and 2011 PIWT test campaigns for 

uncontaminated fluid applied to the model. Many other runs were performed for contaminated fluids, but 

are not discussed in this report. For each simulated takeoff, the percent lift loss from the clean or dry wing 

was evaluated at αw = 8° as described in Section 3.2. These data are summarized as a function of the test 

temperature in Figure 20. There were five different fluids tested: 

• ABC-S Plus manufactured by Kilfrost, 

• EG106 manufactured by Dow Chemical, 

• AD-49 manufactured by Dow Chemical, 

PIWT Model Angle of Attack, αw, deg.

PIWT Model Lift Coefficient,

Cl

-3.0 -2.0 -1.0 0.0 1.0 2.0 3.0 4.0 5.0 6.0 7.0 8.0 9.0

0.0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

1.6

1.8 Dry Wing

Fluid Only

Fluid and Ice Pellets

4.2% Lift loss

from Dry Wing

8.0% Lift loss

from Dry Wing
NASA/TM—2012-216014 25

• Launch manufactured by Clarient, 

• Max-Flight manufactured by Octagon Process. 

Of these, all are PG fluids except for EG106 which is an EG fluid. The data show the expected variation 

of PIWT model lift loss measured at αw = 8° with test temperature. In general, the percent lift loss 

increases with decreasing temperature. It was noted that EG fluids tend to flow off more readily than PG 

fluids, explaining the lower lift losses in Figure 20 for the EG106 fluid. It is important to note that all of 

these fluids are qualified fluids and as such, have passed the AAT. The variation in percent lift loss 

characteristics is due to the differences in the various fluids and the resulting flowoff behavior.

3.4 Summary of AAT Results

Samples of the fluids tested during the 2010 and 2011 PIWT test campaigns were sent to AMIL at 

University of Quebec at Chicoutimi. These fluids were tested according to the procedures for the highspeed ramp of the AAT. This procedure is described in detail in Reference 12 and was summarized in 

Section 2.5. The test temperature range was selected to correspond to the range used in the PIWT test 

campaigns that was illustrated in Figure 20. The AAT test results for these fluid samples is shown in 

Figure 21. The upper limit of boundary-layer displacement thickness at –20 °C (δ* = 9.74 mm) was 

determined according to Reference 12. The plot shows that all of the fluids tested passed the AAT since

all of the δ* data lie below the upper limit of 9.74 mm. The fluid samples tested were taken from the same 

batch used for the PIWT testing thereby eliminating any potential variations due to the manufacturing 

tolerances of the fluids. In the case of the Launch fluid, the same batch was used for both 2010 and 2011 

test campaigns. The data presented in Figure 20 and Figure 21 were subsequently used to relate the PIWT 

model lift loss to the lift loss on a B737-200ADV airplane using the latter’s relationship to the AAT. 

Figure 20.—Summary of uncontaminated fluid tests for PIWT test campaigns in 2010 and 2011.

⊗

⊗

⊗

⊗

⊗

⊗ ⊗

Test Temperature in PIWT, °C

PIWT Model Lift Loss at

α

w = 8 deg., %

-25.0 -20.0 -15.0 -10.0 -5.0 0.0

0.0

1.0

2.0

3.0

4.0

5.0

6.0

7.0

8.0

9.0

10.0

ABC-S Plus 2010

ABC-S Plus 2011

EG106 2010

EG106 2011

Launch 2010

Launch 2011

AD-49 2011

⊗ Max-Flight 2011
NASA/TM—2012-216014 26

Figure 21.—Summary of AAT results for fluid samples used in the 2010 and 2011 PIWT test campaigns.

3.5 Establishment of a PIWT Model Lift Loss Criterion Via the AAT

As described in Section 2.5 (cf. Figure 15), the results of the AAT relate directly to the maximum lift 

loss on the B737-200ADV airplane. It is therefore possible to use the results of the AAT to relate the lift 

loss measured on the PIWT model to the maximum lift loss on the B737-200ADV airplane. The first step 

in the process was to associate values for displacement thickness with the temperatures used in the PIWT 

testing by two-point interpolation or extrapolation employing the temperatures and displacement 

thicknesses from the AAT results (cf. Figure 21). This step accounts for the fact that the test temperature 

of the runs conducted in the PIWT (cf. Figure 20) were not identical to the test temperatures for the AAT

results (Figure 21). For example, three AAT runs were performed with the EG106 2011 fluid batch at

temperatures of –9.9, –15.7 and –23.0 °C (cf. Figure 21). These temperatures and measured values of δ* 

are shown in Table V. The PIWT fluid tests using the same EG106 2011 batch were conducted at 

temperatures of –10.4, –11.8, –13.4, –15.0 and –23.3 °C. Two-point linear interpolation was performed to 

associate the δ* values with the PIWT temperatures. Since the PIWT test temperature of –23.3 °C was 

slightly lower than the AAT temperature value of –23.0 °C, a linear extrapolation was performed for this 

case using the nearest two temperatures (–15.7 and –23.0 °C). The PIWT temperatures and associated

displacement thickness values are listed in Table VI. The same process was followed for the other fluids. 

The associated δ* values are plotted against the PIWT test temperatures in Figure 22. 

⊗

⊗ ⊗

Test Temperature, °C

Boundary-Layer Displacement Thickness

δ*, mm

-25.0 -20.0 -15.0 -10.0 -5.0 0.0

0.0

1.0

2.0

3.0

4.0

5.0

6.0

7.0

8.0

9.0

10.0

11.0

12.0

ABC-S Plus 2010

ABC-S Plus 2011

EG106 2010

EG106 2011

Launch 2010/2011

AD-49 2011

Max-Flight 2011

δ* limit for -20 °C

δ* dry plate

⊗
NASA/TM—2012-216014 27

TABLE V.—AAT RESULTS FOR 

EG106 2011 FLUID BATCH

AAT test temperature, 

°C

Displacement thickness (δ*), 

mm

–9.9 4.91

–15.7 6.52

–23.0 7.10

TABLE VI.—DISPLACEMENT THICKNESS VALUES 

CORRECTED FOR TEMPERATURE FOR PIWT 

TEST OF EG106 2011 FLUID BATCH

PIWT test temperature,

°C

Interpolated displacement

thickness (δ*), 

mm

–10.4 5.05

–11.8 5.44

–13.4 5.88

–15.0 6.33

–23.3 7.12

Hill and Zierten (Ref. 13) suggested that there is a linear relationship between the maximum lift loss 

on the B737-200ADV airplane and the AAT boundary-layer displacement thickness. This line is shown in 

Figure 15 and is given by: 

∆CL,max = 5.24% � δfluid

∗ −δdry∗

δl @−20 °C ∗ −δdry

∗ � (7) 

This equation was used to calculate maximum lift loss on the B737-200ADV airplane using the data in 

Figure 22. As shown in Figure 22, all of the displacement thickness values are known: 

• δ*fluid is given for each fluid batch as a function of test temperature; 

• δ*dry is given from the AAT data, = 2.66 mm; 

• δ*l @ –20 °C is given from the AAT data, = 9.74 mm. 

Using these data, all of the corresponding values for the B737-200ADV maximum lift loss were 

calculated for each fluid tested in the PIWT at each temperature and the resulting data are plotted in 

Figure 23. As noted previously in this section, all of these fluids passed the AAT so all maximum lift loss 

values lie below the 5.24 percent limit as indicated in Figure 23. The form of the data in Figure 23 is 

similar to that in Figure 22 because of the linear relationship given in Equation (7). 

This analysis has generated the data required to relate the lift losses measured on the thin, highperformance wing model in the PIWT to the B737-200ADV maximum lift loss implied by the AAT. The 

PIWT model lift losses measured at αw = 8° were plotted against test temperature in Figure 20 for tests with 

the fluids shown in the plot legend. The B737-200ADV airplane maximum lift losses were obtained from 

the AAT for the same fluids at the same test temperatures as shown in Figure 23. Therefore, the common 

independent variable of test temperature can be eliminated between Figure 20 and Figure 23 to yield a 

correlation between the B737-200ADV maximum lift loss and the PIWT model lift loss measured at 

αw = 8°. This correlation is shown in Figure 24 and reflects the appropriate trend of increasing B737-

200ADV maximum lift loss for increasing PIWT model lift loss measured at αw = 8°. This plot is 

analogous to that shown in Figure 13 relating B737-200ADV three-dimensional model maximum lift losses 

to the lift losses on the corresponding two-dimensional model at αw = 8°. In both plots (Figure 13 and 

Figure 24) there is considerable scatter in the data that is indicative of the differences in the fluid flowoff 
NASA/TM—2012-216014 28

Figure 22.—Boundary-layer displacement thickness values associated with PIWT test temperatures. 

Figure 23.—B737-200ADV maximum lift loss for the PIWT test temperatures calculated using AAT data 

shown in Figure 22 using Equation (7).

⊗ ⊗ ⊗ ⊗ ⊗⊗ ⊗

Test Temperature in PIWT, °C

Boundary-Layer Displacement Thickness,

δ*, mm

-25.0 -20.0 -15.0 -10.0 -5.0 0.0

0.0

1.0

2.0

3.0

4.0

5.0

6.0

7.0

8.0

9.0

10.0

11.0

12.0

ABC-S Plus 2010

ABC-S Plus 2011

EG106 2010

EG106 2011

Launch 2010

Launch 2011

AD-49 2011

Max-Flight 2011

δ* limit for -20 °C

δ* dry plate

⊗

⊗ ⊗ ⊗ ⊗ ⊗⊗ ⊗

Test Temperature, °C

B737-200ADV Max. Lift Loss,

∆

CL,max , %

-25.0 -20.0 -15.0 -10.0 -5.0 0.0

0.0

1.0

2.0

3.0

4.0

5.0

6.0

ABC-S Plus 2010

ABC-S Plus 2011

EG106 2010

EG106 2011

Launch 2010

Launch 2011

AD-49 2011

Max-Flight 2011

5.24% Limit

⊗
NASA/TM—2012-216014 29

Figure 24.—Correlation between percent lift loss on the PIWT model at αw = 8° and percent maximum lift 

loss on the three-dimensional B737-200ADV airplane for the uncontaminated fluid tests. 

characteristics and resulting performance penalties. A linear regression curve fit was performed on the 

data in Figure 24. The regression was forced through the origin since this corresponded to the dry wing 

for each configuration. The percent lift loss on the dry wing was 0, by definition. The regression line had 

a slope = 0.72 whereas the slope in Figure 13 was 1.04. The most likely explanation for the difference in 

these slopes is that the thin, high-performance wing section used in the PIWT test had significantly 

different aerodynamic characteristics than the “critical” section of the B737-200ADV wing. There are 

other differences as well, such as the speed at model rotation, between the two tests. Therefore some 

difference in these slopes is expected. The data indicate that a ΔCL,max of 5.24 percent on B737-200ADV 

airplane is equivalent to a lift loss of 7.3 percent on the PIWT model measured at αw = 8°. It is therefore 

suggested that an appropriate lift loss threshold for the tests conducted on the PIWT model is 7.3 percent

measured at αw = 8° since this is equivalent to a ΔCL,max of 5.24 percent on B737-200ADV airplane. 

It is clear from inspection of Figure 24 that some means of accounting for the scatter in the data is 

needed. Furthermore, establishing an absolute lift loss threshold of 7.3 percent neglects the three data 

points in Figure 24 that have a PIWT model lift loss measured at αw = 8° greater than 7.3 percent. These 

cases should not fail the criterion because they are qualified fluids having passed the AAT. Chapra and 

Canale (Ref. 19) define the Standard Error of the Estimate (SEE) as a means of quantifying the data 

spread about the linear regression line and is given by: 

SEE = �∑ (yi−0.72xi) N 2 i=1

N−2 (8) 

The SEE is a measure of the “goodness of fit” based upon the sum of the residuals about the regression 

line. For the data in Figure 24, SEE = 0.67 and has “units” of the y-variable—B737-200ADV airplane 

percent maximum lift loss. Plotted in Figure 24 are lines representing ±2*SEE deviations from the linear 

regression line and all but two of the data points fall between these lines. In addition, the upper bound 

(+2*SEE) captures the three PIWT data points with lift loss greater than 7.3 percent but have still passed 

⊗

⊗ ⊗

⊗

⊗⊗ ⊗

PIWT Model Lift Loss at αw = 8 deg., %

B737-200ADV Max. Lift Loss,

∆

CL,max , %

0.0 1.0 2.0 3.0 4.0 5.0 6.0 7.0 8.0 9.0 10.0

0.0

1.0

2.0

3.0

4.0

5.0

6.0 ABC-S Plus 2010

ABC-S Plus 2011

EG106 2010

EG106 2011

Launch 2010

Launch 2011

AD-49 2011

Max-Flight 2011

5.24% Limit

Linear Fit y = 0.72x

Linear Fit ± 2*SEE

⊗

5.4 7.3 9.2
NASA/TM—2012-216014 30

the AAT. Therefore, it seems prudent to establish a lift loss threshold up to 9.2 percent for the PIWT 

model lift loss at αw = 8°. Since the evaluation of the fluid performance with ice-pellet contamination is 

also based upon visual observations of flowoff and failure, specifying a range of lift loss can help identify 

cases that may need more careful evaluation of the visual observations. For example, if the PIWT 

measured lift loss at αw = 8° is greater than 7.3 percent, but less than 9.2 percent, then the fluid 

performance may still be acceptable based upon visual observations. However, if the PIWT measured lift 

loss at αw = 8° is greater than 9.2 percent, this should be grounds for rejection since the upper limit has 

been surpassed. Additional data may result in future adjustments to the proposed criterion. 

4.0 Comments on PIWT Model Tests and Results

As stated in Section 1.0, the dissemination of the PIWT test results with the thin, high-performance 

wing model and the subsequent changes to the ice-pellet allowance time tables have led to concerns 

regarding the fidelity and applicability of the measured lift degradation. Given the potential operational 

safety implications associated with the measured performance degradations, it is even more important that 

these concerns be addressed appropriately. The essence of these concerns is understanding how the 

performance degradations measured on the model in the NRC PIWT testing relate to an actual threedimensional, full-scale airplane configuration. The review of the research basis of the AAT and the 

subsequent analysis in Section 3.0 was conducted to address the following specific concerns: 

• A two-dimensional model with endplates was used instead of a three-dimensional geometry more 

representative of an airplane wing. 

• Lift degradations were measured at αw = 8° on the two-dimensional model instead of at stall.

• Airplanes undergoing take-off accelerations and rotation experience ground effects that were not 

simulated in the PIWT testing. 

• The chord Reynolds number was small relative to most airplane wing sections.

• The 6-ft (1.8 m) chord length model is small relative to most airplane wing sections.

• The large model size and proximity of the wind-tunnel walls resulted in uncorrectable effects on 

the aerodynamic data and performance.

• The rotation speed used in the PIWT tests was low relative to typical airplane rotation speeds.

Interestingly, all but the last two of these concerns were also raised in the development of the AAT 

and were addressed by Hill and Zierten (Ref. 13) and Runyan et al. (Ref. 16). The experimental methods 

used in the aerodynamic evaluations of fluids in the PIWT closely follow the methods used in the twodimensional model testing during the development of the AAT. Hill and Zierten (Ref. 13) and Runyan et 

al. (Ref. 16) showed the usefulness and applicability of two-dimensional model testing. In both research 

efforts, lift losses on the two-dimensional model were evaluated at an angle of attack well below stall.

The PIWT model lift data were analyzed at αw = 8°. At this low angle of attack, the flow on the model 

was two dimensional based upon observations of the fluid flow. So analyzing the lift data at αw = 8°

minimized any three-dimensional effects and wall effects that may increase in magnitude with increasing

angle of attack above 8°. In both research efforts, the wind-tunnel testing was performed without a ground 

plane (for the two-dimensional model). In the AAT research, the two-dimensional model was very small 

(1.5 ft (0.305 m) chord) relative to the full-scale wing. The small size of the wing meant that the chord 

Reynolds number (≈1.5×106 at rotation) was also very low. The PIWT model was larger (6 ft chord) and 

therefore higher Reynolds numbers (≈ 7.0×106 at rotation) were obtained in the wind tunnel. The effect of 

Reynolds number on aerodynamic performance of airfoils is usually insignificant above 6.0×106

. This 

“critical” Reynolds number is much lower with any type of leading edge contamination on the wing 

(Refs. 20 to 23). The AAT research indicated model scale and Reynolds number effects were 

negligible—down to the small values used in that work. Therefore, the present PIWT model scale and 

Reynolds number should not introduce any additional concerns. 
NASA/TM—2012-216014 31

There is little reason for concern about the model size and wind-tunnel walls resulting in large and 

uncorrectable solid and wake blockage effects that would adversely affect the aerodynamic data. The 

model chord to test-section height ratio was 0.37, which is less than 0.40—the value recommended as the 

reasonable upper limit for two-dimensional model testing for accurate values of maximum lift (Ref. 24).

Since the model was not oversized relative to the tunnel height, the effects of the walls should be

minimal. These effects should be similar for all of the test runs and therefore should not affect the relative 

comparisons. Furthermore, wind-tunnel wall effects become more significant at higher angles of attack 

and lift coefficient approaching maximum lift. So measuring the PIWT model lift loss at αw = 8°, well 

below maximum lift, also mitigated these effects. 

There are some differences between the B737-200ADV research effort used to develop the AAT and 

the present PIWT model research effort. However, these differences in the experimental methods, such as 

model size, geometry, installation and Reynolds number, should be accounted for in the correlation 

presented in Figure 24. This discussion has shown that the concerns listed at the beginning of this section 

do not compromise the applicability of the AAT scaling methodology for PIWT model lift loss. The 

PIWT testing was carried out in a manner consistent with the methods used in the AAT research and with 

standard wind-tunnel testing methods. The possible exception is the lower rotation speed, addressed in the 

following paragraph. 

In the PIWT model tests, the speed at model rotation was limited to 100 kn (51.4 m/s) by the 

capabilities of the wind-tunnel facility. In the “high-speed ramp” procedures currently used for the AAT, 

the simulated rotation speed is 127 kn (65 m/s) (Ref. 12). This value is very similar to the rotation speed 

of 120 to 125 KEAS (62 to 64 m/s) that was used in the flight test trials and wind-tunnel testing on the 

B737-200ADV for the development of the AAT (Refs. 13 and 16). A similar rotation speed range was 

used in the Falcon 20 trials described in Section 1.0. The effect of rotation speed on lift coefficient was 

shown in Figure 5 and is significant. Higher rotation speeds resulted in higher lift coefficient, which in 

turn means lower percent lift loss due to the fluid. The relationship between rotation speed and lift 

coefficient appears to be linear over the range shown in Figure 5. If this relationship is linear, then the 

primary effect of low rotation speed on the PIWT aerodynamic data may already be accounted for in the 

“scaling” analysis described in the previous section. The low rotation speed in the PIWT tests could partly 

explain the lower slope of the regression line plotted in Figure 24. However, more data may be required to 

completely ameliorate this concern. 

There are additional concerns not included in the list at the beginning of this section. For example, the 

AAT was developed prior to the current widespread use of Type IV anti-icing fluids, such as those used 

in the PIWT tests. If adjustment to the AAT acceptance criteria were required for Type IV fluids, then this 

may affect the proposed “scaling” methodology applied to the PIWT data. Beisswenger et al. (Ref.25 ) 

performed a series of flat-plate fluid tests comparing flowoff characteristics that showed similar behavior 

between Type II and Type IV fluids. The authors are unaware of any corresponding aerodynamic 

evaluation that would help confirm similarities between Type II and Type IV fluid effects and further 

research may be required. 

Another concern that may be introduced by the proposed “scaling” methodology relative to the 

development of the ice-pellet allowance times is that the visual observations of the fluid flowoff and 

associated evaluation are still based upon the aerodynamics of the thin, high-performance wing while the 

lift loss criterion has been scaled to the effect on the full-scale B737-200ADV airplane. This represents 

some type of mixed evaluation that may or may not affect the development of the allowance times. For 

example, a key difference in the model geometries is that the B737-200ADV model had a leading-edge 

slat while the PIWT model had a hard leading edge. Hill and Zierten (Ref. 13) noted this “configuration 

dependency” in their work. It is unclear that similar ice-pellet contamination testing on the B737-

200ADV model would lead to similar allowance times. Finally, it may be worthwhile revisiting the fluids 

research on commuter aircraft that was conducted shortly after the B737-200ADV IRT test campaigns, 

described in References 26 and 27. This may help to further mitigate concerns about the present PIWT 

model tests and results due to a higher degree of similarity in the model configurations and rotation 

speeds. 
NASA/TM—2012-216014 32

5.0 Summary and Conclusions

The FAA has worked with Transport Canada and others to develop allowance times for aircraft 

operations in ice-pellet precipitation. Wind-tunnel testing has been carried out to better understand the 

flowoff characteristics and resulting aerodynamic effects of anti-icing fluids contaminated with ice 

pellets. This testing has taken place at the National Research Council of Canada’s Propulsion and Icing 

Wind Tunnel (PIWT) using a nominally two-dimensional wing with a thin, high-performance section and 

single-element slotted flap. The model, with either uncontaminated anti-icing fluid or fluid contaminated 

with simulated ice-pellet precipitation, was subjected to a simulated takeoff through rotation. Evaluation

of the fluid performance was determined from visual observations of the flowoff and evaluation of the 

model lift degradation. The percent lift loss on the PIWT model was calculated at 8° angle of attack and 

used as one of the evaluation criteria in determining the allowance times in recent testing. The 

dissemination of the wind-tunnel testing results and subsequent changes to the ice-pellet allowance time 

tables led to concerns regarding the fidelity and applicability of the measured lift degradation in fluid 

testing at the NRC PIWT. Given potential operational safety implications associated with the measured 

performance degradations, it was important that these concerns be addressed. This report addresses these 

concerns and shows that the lift loss in the PIWT can be related to the loss in maximum lift of a Boeing 

737-200ADV airplane through the Aerodynamic Acceptance Test (AAT) performed for fluids 

qualification. 

The AAT was developed as a simple and relatively low-cost method to evaluate deicing and antiicing fluids. Fluid samples are tested on a small, horizontal, flat plate at operational air temperatures and 

simulated takeoff speeds. The measured boundary-layer displacement thickness due to the fluid was 

shown to be directly proportional to the loss in maximum lift on a B737-200ADV airplane. Fluids with 

AAT displacement thickness values corresponding to less than 5.24 percent maximum lift loss are 

considered acceptable. Fluid samples used in the PIWT model testing were subjected to the AAT to 

develop the relationship for those samples between displacement thickness and B737-200ADV maximum 

lift loss. Samples of these same fluids were tested at the same temperatures on the thin, high-performance 

wing model in the PIWT and lift losses were determined at 8° angle of attack. This permitted the 

development of a correlation between the PIWT model lift loss at 8° angle of attack and lift loss at

maximum lift on the B737-200ADV airplane. 

The analysis presented in this report shows that a maximum lift loss of 5.24 percent on the B737-

200ADV airplane corresponds to a lift loss of 7.3 percent for the thin, high-performance wing at αw = 8°

in the PIWT. There is significant scatter in the data used to develop the correlation related to varying 

effects of the various anti-icing fluids that were tested and other factors. A statistical analysis indicated 

that the upper limit of lift loss on the PIWT model was 9.2 percent. Therefore, a range of lift loss from 7.3 

to 9.2 percent on the thin, high-performance wing at αw = 8° in the PIWT suggests that extra scrutiny of 

the visual observations is required in evaluating fluid performance with contamination and establishing 

the allowance times. For example, if the PIWT model lift loss at αw = 8° is greater than 7.3 percent, but 

less than 9.2 percent, then the fluid performance may still be acceptable based upon visual observations.

However, if the PIWT model measured lift loss at αw = 8° is greater than 9.2 percent, this should be 

grounds for rejection since the upper limit has been surpassed.

This review of the research basis of the AAT and the subsequent PIWT model analysis has shown 

that any differences such as model size, geometry, installation and Reynolds number, are accounted for in 

the correlation presented in Figure 24. Furthermore, this report shows that concerns about the thin, highperformance wing model testing in the PIWT do not compromise the applicability of the AAT scaling 

methodology for PIWT model lift loss determined at 8° angle of attack. The PIWT testing was carried out 

in a manner that was consistent with the methods used in the AAT research and with standard windtunnel testing methods. There are, however, some remaining concerns such as rotation speed, 

applicability to Type IV fluids and model configuration that could be addressed in future research.
NASA/TM—2012-216014 33 

References

1. FAA Notice N 8000.308, “Reevaluation of Deicing/Anti-icing Programs,” Oct. 5, 2005. 

2. FAA Notice N 8000.309, “Dispatching During Precipitation Conditions of Ice Pellets, Snow Pellets, 

or Other Icing Events for which no Hold Over Times Exist,” Oct. 5, 2005. 

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Operations in Ice Pellets without Deice/Anti-Ice Fluids,” Oct. 31, 2005. 

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Snow Conditions and Evaluating Pilot Assessment of Precipitation Intensity,” Sept. 20, 2007. 

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http://www.faa.gov/other_visit/aviation_industry/airline_operators/airline_safety/deicing Accessed

Sept. 7, 2011. 

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Aerospace Sciences, Vol. 37, No. 8, Nov. 2001. 
NASA/TM—2012-216014 34

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Aerospace Sciences, Vol. 41, No. 5, Jul. 2005, pp. 323-362. 

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Consideration of the Effects of Compressibility,” NACA Report 782, 1944. 

25. Beisswenger, A., Fortin, G., Laforte, J.-L., “Investigation of Type II and Type IV Aircraft Ground 

Anti-Icing Fluid Aerodynamic Certification Standards,” FAA Report DOT/FAA/AR-03/55, July 

2003. 

26. Ellis, N.D., Nettleton, T.R., and Eggleston, B., Effects of Anti-icing/De-icing Fluids on the Take-off 

Performance of Commuter Aircraft,” Report DHC-TDC-90-1, Transport Canada DeHavilland, Mar. 

1991. 

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