NASA/TM--2001-210907 AIAA-97-0051

Validation of NASA Thermal Ice Protection

Computer Codes

Part 3: The Validation of Antice

Kamel M. A1-Khalil and Charles Horvath

Cox & Company, Inc., New York, New York

Dean R. Miller

Glenn Research Center, Cleveland, Ohio

William B. Wright

NYMA, Inc., Brook Park, Ohio

May 2001

https://ntrs.nasa.gov/search.jsp?R=20010068034 2018-05-14T14:49:22+00:00Z
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NASA/TM--2001-210907 AIAA-9?-0051

Validation of NASA Thermal

Computer Codes

Part 3: The Validation of Antice

Ice Prote, ction

Kamel M. A1-Khalil and Charles Horvath

Cox & Company, Inc., New York, New York

Dean R. Miller

Glenn Research Center, Cleveland, Ohio

William B. Wright

NYMA, Inc., Brook Park, Ohio

Prepared for the

35th Aerospace Sciences Meeting and Exhibit

sponsored by the American Institute of Aeronautics and Astronautics

Reno, Nevada, January 6-7, 1997

National Aeronautics and

Space Administration

Glenn Research Center

May 2001
Acknowledgments

This work was sponsored by the Icing Technology Branch at the NASA Glenn Research Center, CleveIand, Ohio.

NASA Center for Aerospace Information

7121 Standard Drive

Hanover, MD 21076

Available from

National Technical Information Service

5285 Port Royal Road

Springfield, VA 22100

Available electronically at http://gltrs.grc.nasa.gov/GLTRS
VALIDATION OF NASA THERMAL ICE PROTECTION COMPUTER CODES

PART 3: THE VALIDATION OF ANTICE

Kamel M. AI-Khalil*, Charles Horvath**

Cox & Company, Inc.

200 Varick Street

New York, New York 10014

Dean R. Milier t

National Aeronautics and Space Administration

Glenn Research Center

21000 Brookpark Road

Cleveland, Ohio 44135

William B. Wright t

NYMA, Inc.

2001 Aerospace Parkway

Brook Park, Ohio 44142

ABSTRACT

An experimental program was generated by the Icing Technology Branch at NASA

Glenn Research Center to validate two ice protection simulation codes: LEWlCE/'l'hermal

for transient electrothermal de-icing and anti-icing simulations, and ANTICE for steady state

hot gas and electrothermal anti-icing simulations. An electrothermal ice protection system

was designed and constructed integral to a 36 inch chord NACAO012 airfoil. The model was

fully instrumented with thermocouples, RTD's, and heat flux gages. Tests were conducted

at several icing environmental conditions during a two week period at the NASA Glenn Icing

Research Tunnel Experimental results of running-wet and evaporative cases were

compared to the ANTICE computer code predictions and are presented in this paper.

Nomenclature mevap = rate of surface water evaporation

(lbm/hr._)

AOA =angle of attack of body with the m_,,p = rate of water droplets impingement

freestream air (degrees) (Ibm/hr._)

Cpw = specific heat (Btu/Ibm.°F) q = heater power density (W/in 2)

h = heat transfer coefficient between the dry qoonv = convective heat loss (W/in 2)

surface and the ambient air (Btu/hr._.°F) qevap = evaporative heat loss (W/in 2)

hf = film coefficient between the wet surface qKE = kinetic heat gain from droplets impact on

and the ambient air (Btu/hr._.°F) surface (W/in 2)

K = thermal conductivity (Btu/hr.ft.°F) qse,s = sensible heat required to warm the

LWC = liquid water content in the freestream impinging water droplets (W/in 2)

(g/m 3) S = surface distance from stagnation, positive

Lv = water latent heat of vaporization on the upper surface (in)

(Btu/Ibm) Taw = adiabatic wall temperature (°F)

MVD = mean volume droplet diameter in cloud T s = skin temperature (°F)

(_m) Trot = total ambient temperature (°F)

m = surface runback mass flow rate per unit Too = static ambient temperature (oF)

span distance (lbm/ff.hr) V = freestream velocity (mph)

= droplets collection efficiency

*Engineedng Scientist, Member AIAA 13 = mass density (Ibm/fta)

**De-icing Systems Engineer

*Aerospace Engineer, Member AIAA

tAerospace Engineer, Member AIAA

NASA/TM--2_I-210_7 1
I. Introduction

xtensive testing is normally required in the

design of thermal ice protection systems. The

cost involved can be somewhat prohibitive. Some

designers reduce costs by relying upon computer

simulations to get through the preliminary design

stage. The total cost and development period

involved can further be reduced if the simulation

codes are proven and reliable. Additionally, the

simulation codes may be utilized in the certification

process, in conjunction with flight/tunnel testing,

once approved by the FAA.

An experimental program was initiated and

sponsored by the Icing Technology Branch at the

NASA Glenn Research Center (GRC) in an effort to

validate two thermal ice protection simulation

codes: LEWICE/Thermal simulates the transient

response of electrothermal de-icing and anti-icing

systems, and ANTICE simulates the steady state

response of hot gas and electrothermal anti-icing

systems. The effort was a partnership between

NASA GRC and five US industry members.

Participants from industry assumed the role of

technical advisors. Additionally, Cox & Company,

Inc. provided the test article, test monitoring and

control software/hardware, and engineering test

support.

An overview of the program is discussed by

Miller, 1 and the validation of the LEWlCE/Thermal

code is presented by Wright. =This particular paper

is dedicated to the validation of the ANTICE code

which is a two dimensional computer simulation of

a composite structure with embedded heaters. It

also simulates a hot gas anti-icing system. The

code has a rivulet model to simulate the flow of

runback water on the surface. Experimental antiicing data will be presented, and comparison will be

made to the ANTICE code predictions.

II. Analytical Model

The computer code is based on an analytical

thermodynamic model that was described in earlier

studies by AI-Khalil. 3'4The vadous modes of heat

transfer occurring within an anti-iced aircraft surface

and the collected water, referred to as "runback"

when it flows on the surface, have been modeled.

The mathematical formulation is based on 2-D

steady-state heat transfer model along a surface

streamline.

The temperature is assumed to vary in the

streamwise direction and across the structure

thickness (transverse direction). Spanwise

variations are much smaller than those in the

chordwise direction. The runback water flow is also

assumed, as proven in several tunnel tests, to be in

the streamwise direction on the aircraft surface.

The aircraft structure may be represented by a

number of composite layers. Any single layer can

be composed of several heater zones, each with a

specified power density. The mathematical

formulation of the energy balance equation in the

structure is based on the control volume approach.

The structure is subdivided into a number of grids in

the streamwise and transverse directions. An

equation is then written for each node in the

discretized region using the laws of energy

conservation (and mass conservation on the

impinging supercooled water droplets).

Anti-icing of the aircraft surface may also be

accomplished using a hot air system. This is

currently modeled such that hot air of a given mass

flowrate and an initial temperature is supplied to the

inner surface in the vicinity of the stagnation line,

and flows back into the cowl in the direction of

surface streamlines. A user can also specify a

distribution of variable heat flux at the surface to

obtain an approximate system behavior as a first

step in the design and/or analysis process (i.e.,

evaporative, running wet, or freezing runback).

Also modeled is the runback flow behavior on

the surface: continuous surface coverage at the

direct impingement regions, versus rivulet structure

flow downstream. Rivulets are individual narrow

streams of water. A stability analysis was performed

for the prediction of rivulets? '4

ANTICE requires the following information:

(1) Surface pressure coefficient distribution along

a streamline.

(2) The distance between two adjacent surface

streamlines representing a zone of interest in

the analysis. This is an option for 3-D flows

such as on swept wings or inlet nacelles.

(3) Collection efficiency distribution at the surface.

The above information may be obtained from flow

and trajectory codes such as NASA's LEWlCE ice

accretion code. In the current analytical predictions,

LEWICE 5 vl.7 was used to obtain the latter

information. Details on the numerical procedure are

presented in Reference [4].

NASA/TM--2001-210907 2
III. Experimental setup

II1.1 Test Article

A NACA0012 airfoil, 6 ft span and 36" chord, was

fitted with an electro-thermal ice protection system

at the leading edge (LE) and mounted vertically in

the NASA GRC Icing Research Tunnel (IRT). The

removable LE was 10" long, and the wooden

afterbody was 26" long. The ice protection system

consisted of seven heater bands, each 18 inches

long in the spanwise direction, individually

controlled by the Cox Thermal Test Management

System (TTMS). For redundancy, a spanwise image

mirror of the seven heaters was constructed around

the center line of the airfoil. The total heated

spanwise length was 3 ft, and the total heated wrap

distance in the chordwise direction was 7.75 inches.

A guard heater (8" span x 12" wrap) was

installed on each end of the heated area along the

airfoil span. The guard heaters were used to

maintain as much of the airfoil clean of ice to

prevent flow disturbances from affecting the

measurement locations near the centerline. The

total number of heated channels was 18.

Figure 1 shows a cross-section of the LE where

zone A is 0.75" wide, zones B, C, D, and E are 1"

wide, and finally zone F and G are 1.5" wide. Zones

B, D, and F are on the IRT control room side.

During the manufacturing process, the heater lay-up

was accidentally shifted towards zones C, D, and G.

The shift was estimated at about 0.19" during the

tunnel tests when different amounts of surface

runback were observed on each side of the airfoil.

This heater shift was modeled in the computer code

for a proper comparison between the experimental

data and the analytical predictions.

Figure 2 illustrates the material composition of

the LE structure in the heated zones, and the

estimated/known material properties are given in

Table I. This particular type of construction was

required to meet the goals set forth for the

validation program.

111.2 Test Instrumentation

The model was heavily instrumented with type-T

thermocouples, heat flux gages, and Resistance

Temperature Detectors (RTD). A total of 76 sensors

was used. The locations of several of the sensors

are illustrated in Reference [1].

The development of the Cox TTMS began in

the early 1980's, and has been used to support the

design of ice protection systems for the F-14, OV-1,

AH-64, B-1B, and SAAB 340 aircraft, among others.

It performs two main functions: data acquisition and

closed-loop control. It consists of data acquisition

boards to measure analog input signals which are

sent to a personal computer via data bus. These

signals are converted to digital representation of

physical measurements (temperature, heat flux,

resistance) with a custom written software. The

software then sends signals back to the power

module in the ]-I'MS to control the amount of power

supplied to individual heater zones.

A user can request the TTMS to maintain a

particular sensor at a desired temperature (e.g., to

maintain the thermocouple temperature feedback

on the metal skin constant). The TTMS will then

control in a closed-loop and resolve the power

density required for each heater to yield the desired

temperature. Conversely, the user can simply set

the power density to each heater independent of

sensors temperature feedback (TTMS will

disconnect the power in case the temperature

exceeds a preset value).

Additional equipment was used to monitor the

test. Video recorders were used throughout the

tests to monitor surface runback and the ice growth.

Still photographs were taken with 35 mm cameras

to document ice shapes and residual ice following

each run. An infrared camera setup was utilized to

monitor the airfoil LE surface temperature from the

IRT control room. A portion of the LE was coated

with a thin black paint to increase the surface

emissivity to near 0.96 (Bare metal normally has

very low surface emissivity).

111.3 Test Conditions

Anti-icing and de-icing runs were performed at

two different airspeeds, several airfoil AOA's, and

various meteorological conditions. These icing

conditions are summarized in Table II. This paper is

dedicated to the validation of the ANTICE program.

Therefore, only anti-icing test runs will be discussed

(evaporative and running-wet). For simplicity, and to

focus on the physical analysis of the icing process

with fewer experimental error uncertainties, only

cases with zero degrees angle of attack will be

considered in the comparison process. The large

experimental database generated will be utilized to

continue the validation process further.

NASA/TM--2001-210907 3
IV. Energy Balance Analysis and Discussions

Simulation of the icing process of aircraft surfaces

is a challenging task. Several uncertainties exist as

a result of the large number of variables involved in

this problem. In order to understand the differences

between the experimental data and the theoretical

predictions, the energy balance terms will be

illustrated. The various modes of heat transfer will

be illustrated to study the contribution of each on

the total power required for a running-wet system

and an evaporative system.

The energy required to heat the aircraft skin to

a specified temperature may be expressed in

general terms as:

q = qco°v+ qevap + qsens + qKE (1)

where the individual components are written as,

qcon_= h (Ts-Taw) (2)

qevap = mevap Lv (3)

qseo,= m_mpCpw(T,-To_)

qKE= mimpV2/(2 go)

(4)

(5)

The runback heat of enthalpy was neglected in

Eq. (1) for simplicity in illustrating the point. Besides,

this term is negligible when the surface is controlled

to a specified temperature in which case the

runback is maintained at a constant temperature.

The rate of water impingement is:

mira p = 13 LWC V (6)

The rate of water evaporation from the surface is a

non-linear function of the surface temperature as

described in Reference [6].

In order to identify the contribution of each term,

two example cases will be considered: a cold

condition at -22 °F, and a warm condition at 20 °F.

These examples will be useful in the analysis of the

ANTICE predictions when compared to the

experimental data in the following section. The

following parameters were selected as

representative values:

h = 40 Btu/hr.ft2.°F

LWC = 0.39 g/m 3

13 = 0.5

V = 200 mph

The heat required to keep the surface at

specific range of temperatures was computed using

Eq. (1). The total power density as well as each

heat component were plotted in Figures 3 and 4 for

the warm and cold conditions, respectively. The

corresponding results expressed in terms of

percentage of the total heat required are illustrated

in Figures 5 and 6.

These plots show the strong dependence of

total power required for evaporative anti-icing

(normally in the range of 100 to 120 °F) on the

evaporation rate which is directly related to droplet

impingement rates. On the other hand, to maintain

the surface just above freezing, for example at

40 °F, the total power required is highly dependent

on the convective heat transfer coefficient and

ambient temperature.

V. Test Results and Comparisons

Several runs at different icing conditions (see

Table I) were executed. Experimental results of

most of the runs were summarized in Table III

where the Run numbers and the corresponding

TTMS file record number are shown. The measured

skin temperatures and the heater power densities

for each zone are displayed. The results and

comparisons to the ANTICE predictions will be

illustrated and discussed for each mode of anti-icing

separately, running-wet and evaporative.

V.1 Running-wet Anti-icing Cases:

In the running-wet conditions, the runback was

maintained at temperatures a few degrees above

freezing within the heated region. Beyond this

region, the runback freezes along a distance that

depends on several variables but mostly on

temperature. Figures 10 and 11 show that effect.

ANTICE normally predicts freezing within about

0.75" length for cold conditions and about 2" for

warm and very wet (high LWC) conditions.

In the validation process, the measured power

densities were then used as inputs to the ANTICE

code. Comparisons between the predicted and

measured skin temperatures will be the basis of

validation in this paper. Few example cases will be

illustrated as representative to the level of

agreement between the theoretical and the

experimental data. It is planned that all test runs will

be summarized and made available on digital

media, such as on a CD-ROM, along with ANTICE

predictions for several runs.

NASA/TM--2001-210907 4
In running-wet operations, the surface runback

mass flow rate rapidly increases as it moves away

from the stagnation line within the direct

impingement region. Since a small percentage is

evaporating, the runback flow rate should reach a

maximum just upstream of the impingement limits,

and start to decrease slowly as it flows to

downstream regions. To illustrate this phenomenon,

consider the case of Run#22B (TTMS#71 as shown

in Table III). This corresponds to icing condition 1 as

illustrated in Table I1.

Figure 12 illustrates the experimental as well as

the ANTICE predicted skin temperature and surface

mass flow rate distributions. The above discussion

of surface mass flow rate is predicted by ANTICE.

The surface temperatures are in excellent

agreement with the measured values. This was not

the case for the aft 2 heaters (D, E, F, and G) when

the measured laminar heat transfer coefficients

shown in Figures 8 and 9 were used. When the

experimental film coefficient for the wet condition

was plotted as shown in Figure 13, it was realized

that the laminar flow is disturbed by the surface

runback water (beads and rivulets) and transition to

turbulent flow must have occurred just past about

1.25 inches. Therefore, the flow was manually

tripped at those locations, and 2 new heat transfer

coefficient distributions (one at 1OOmphand another

at 200mph freestream velocities) were used

thereafter in the remainder of the cases.

Additional running-wet cases illustrated in

Figures 14 through 17 show a very good agreement

with the experimental values. It should be realized

that more studies must be conducted to predict the

transition locations for different geometries at

different airspeeds and surface runback physical

structure and dimensions.

V.2 Evaporative Anti-icing Cases:

Evaporative anti-icing predictions were not as good

as the running-wet cases. As discussed in

Section IV, and as shown in Figures 5 and 6 for

temperatures in excess of 100 °F, approximately

70% of the total power required to achieve a fully

evaporative operation (no runback beyond the

heated zones) is the heat component required to

evaporate surface water as given by Equation (3).

This is a result of

vaporization of water,

This phenomenon is better understood through

an example. Consider Run#22A (TTMS#70).

Figure 18 illustrates the experiment film coefficient

in a wet environment. The average wet film

coefficient should superimpose the dry film or heat

transfer coefficient beyond the direct impingement

zone where the surface is dry. This occur near +/-2

inches from the hilite. Beyond those locations, the

experimental film coefficient increases indicating

that transition to turbulence has occurred. In a

similar manner to the running-wet conditions, the

boundary layer was manually tripped at those

locations to predict the heat transfer coefficient

distribution which was used for the remainder of the

cases at the same freestream velocity.

Figure 19 shows the ANTICE code predictions

for the skin temperature and the runback mass flow

rate. The comparison is acceptable considering the

large surface temperature variations within short

distances as also observed in the IR camera

thermographic data. Surface runback water

disappear near and past the impingement limits

confirming an evaporative system operation.

Figure 20 illustrates a typical behavior seen in

many of the cases. In low airspeeds, 1OOmph here,

a higher surface temperature is required to achieve

full evaporation. The two main elements that

promote evaporation are a higher mass transfer

coefficient which is proportional to the heat transfer

coefficient, and a high surface temperature. This

was confirmed in all the experimental results and

shown here in two examples: Figure 20 for the low

airspeed of 100mph and Figure 21 for the high

airspeed 200mph. Both cases were at the same

ambient temperature (0 °F). At 100mph, it was

noticed that a surface temperature of about 150 °F

was required as opposed to 100 °F in the case of

200mph.

Comparisons of the predicted temperatures to

the experimental measurements in those two

figures is subjective. In many cases, the code

seems to predict the stagnation within about 10 °F

which is acceptable considering the measurement

uncertainties and the number of environmental

variables involved. However, the skin temperature

at zones B&C, adjacent to the hilite heater, are

always underpredicted by ANTICE (by 25 °F or

the high latent heat of more). This may be associated with two reasons:

(1) many of the conditions were near the edge of

NAS A/TM--2001 -210907 5
the spray bar calibration envelope which could

result in smaller LWC than actually predicted by the

calibration standard; (2) the predicted rate of mass

evaporation is always assumed to get 100% of the

latent energy from the skin and absolutely none

from the heated boundary layer. A closer look

should be given to all variables throughout the icing

simulation (experimental and computational).

Vi. Concluding Remarks

An experimental program was generated by the

Icing Technology Branch at NASA Glenn Research

Center to validate two ice protection simulation

codes: LEWlCE/Thermal and ANTICE. This paper

was dedicated to the validation of ANTICE, a steady

state hot gas and electrothermal anti-icing

simulations. An electrothermal ice protection system

was designed and constructed integral to a 36 inch

chord NACA0012 airfoil. The model was fully

instrumented with thermocouples, RTD's, and heat

flux gages. Tests were successfully conducted at

several icing environmental conditions in the NASA

Glenn Icing Research Tunnel (IRT). Experimental

results of running-wet and evaporative cases were

compared to the ANTICE computer code

predictions.

The dry heat transfer coefficients predicted by

LEWlCE, when forced to stay laminar, were in

general good agreement with the measured values.

This distribution was slightly higher in the

experimental cases than in the theoretical

predictions. The difference can be attributed to the

fact that the IRT is not considered as a low

turbulence aerodynamic tunnel and its main

purpose is to simulate icing environments. In the

validation process, the experimental values were

used in the laminar region to isolate the effect of

that variable on the ANTICE temperature

predictions.

The computer code predictions for running-wet

anti-icing conditions were very close to the

experimental measurements. However, this

agreement near the aft zone heaters was realized

when the laminar boundary layer (BL) was forced to

trip at these locations as depicted in the

experimental results of the film coefficient

distribution in wet conditions.

The ANTICE code predictions of the skin

temperatures in the evaporative cases were

acceptable near stagnation, but marginal elsewhere

especially at zones B&C. This is due to the large

effects of the evaporative cooling which is directly

attributed to the high latent heat of vaporization of

water. An accurate measurement of LWC and

verification of the collection efficiency distribution

are required to evaluate any deficiencies in the

standard evaporative terms used in most icing

simulation codes. Nevertheless, ANTICE predicted

no runback existed past the heated zones in most

cases. Interestingly, evaporative anti-icing is more

dependent on the power requirements and not the

standard surface temperature of 120 °F normally

assumed. In low velocity conditions (low h), the

surface reached over 150 °F to achieve fully

evaporative conditions.

One should not expect to always correctly

predict the skin temperature at the aft heater zones

because the difference between having little or no

runback at those regions can be as high as 100 °F.

Heaters at those regions must be designed such

that they protect against the majority of icing

condition requirements without exceeding the

material temperature limits in the dry conditions at

the low aircraft airspeed of its operational envelope.

Future recommended studies:

. Develop a model that will predict the transition

location and the heat transfer coefficients for

different conditions. The BL seems laminar for

a relatively wide portion of the leading edge

(LE) as opposed to what LEWlCE normally

predicts. Experimental results show that the BL

is definitely affected by surface water runback

(beads and rivulets). The transition location is

moved further downstream in evaporative antiicing cases but the transition to turbulent is

milder.

. Revisit the evaporative cooling law as applied

to aircraft icing. Studies of the basic physics

must be conducted to isolate the various modes

of heat transfer, especially the evaporative

mass and energy transfer from the surface.

These studies must also verify the cloud LWC

and the water droplet collection efficiency at the

LE of the model. The effect of heating the

surface and evaporation on the BL must also be

considered.

. Carefully analyze all the records in the relatively

large experimental database, and conduct a

thorough validation process with sensitivity

analysis. This, for example, includes verification

of thermocouple readings with thermographic

data from the Infrared camera.

i

NASA/TM--2001-210907 6
References

_Miller, D.R., Bond, T.H., Sheldon, D.W.,

Wright, W.B., Langhals, T., AI-Khalil, K.M.,

Broughton, H., "Validation of NASA Thermal Ice

Protection Computer Codes: Part 1 - Program

Overview", AIAA Paper 97-0049, Jan 1997.

4AI-Khalil, K.M., Keith, T.G., and De Witt, K.J.,

"Development of an Improved Model for Runback

Water on Aircraft Surfaces," AIAA Paper 92-0042,

Jan. 1992.

5Wright, W.B., "Users Manual for the Improved

NASA Lewis Ice Accretion Code LEWlCE 1.6",

NASA CR-198355, June 1995.

2Wright, W.B., AI-Khalil, K.M., Miller, D.R.,- -

"Validation of NASA Thermal Ice Protection

Computer Codes: Part 2 - The Validation of

LEWICE/ThermaI", AIAA Paper 97-0050, Jan 1997.

3AI-Khalil, K.M., Keith, T.G., and De Witt, K.J.,

"Further Development of an Anti-Icing Runback

Model," AIAA Paper 91-0266, Jan. 1991.

6AI-Khalil, K.M., Keith, T.G., De Witt, K.J.,

Nathman, J.K., Dietrich, D.A., 'q'hermal Analysis of

Engine Inlet Anti-Icing Systems," Journal of

Propulsion and Power, Vol. 6, No. 5, Sept-Oct,

1990, pp. 628-634.

ITEM

Table I: Leading Edge Material Properties

MATERIAL K p Cp

(Btu/hr.ft.°F) (ibrn/ft 3) (Btu/lbm.°F)

1 Heating Element (alloy 90) 23.7 556 0.092

2 Erosion Shield (SS 301 HH) 9.4 501 0.12

3 Elastomer (Cox 4300) 0.148 86.4 0.30 +/-0.03

4 Fibe rglass/Epoxy Composite 0.17 112 0.375

5 Silicone Foam Insulation 0.07 40.5 0.27 +/-0.03

Table Ih Icing Conditions Tested at the NASA GRC IRT

Icing

Condition

2

mtot

(°F)

2O

2O

V

(mph)

100

100

LWC

(g/m 3)

.78

1.1

MVD

(l_m)

20

2O

AOA

(degrees)

0

0

3 20 200 .39 20 0

4 0 100 .78 2O 0

5 0 100 1.1 20 0

9 20

10 20

6 0 .39 20 0

7 20 .55 20 0

8 0 .55 20 0

20 -2

11

.I 200

200

200

100

100

100

200

100

100

100

12

13

2O

.78

.78 20 -4

.9O 0

.39

2.0

.78

,79

-22

15

40

2O

2O

2O

2O

14

15

0

0

-2

-4

NASA/TM--2001-210907 7
Table IIh Summary of Measured Skin Temperatures and Heater Power Densities

(Evaporative and Running-wet Anti-Icing Modes)

Icing AntiCond- Run # Icing

ition (TTMS#) Mode

Zone Zone Zone Zone Zone Zone Zone

A B C D E F G

T, q T, q Ts q I", q T, q T, q T, q

1 22 Evap 117 30 147 21 150 12 160 6.6 162 4.5 157 6.4 156 6.6

(70,71) Wet 38 3.1 40 2.6 40 1.9 40 1.9 39 2.2 40 1.7 38 1.5

2 48 Evap 116 30 158 27 163 19 169 8.2 172 4.3 165 7.6 164 8.8

(75,76) Wet 37 3.2 39 2.6 40 2.3 39 2.0 39 2.3 38 1.7 39 1.7

3 53 Evap 92 24 97 17 100 12 103 6.7 106 3.0 100 6.2 101 5.3

(80,81) Wet 36 4.2 38 3.4 38 3.4 37 3.5 36 3.8 35 2.6 35 2.4

4 35 Evap 112 30 144 22 151 15 154 7.6 157 4.3 150 7.8 149 8.2

(63,64) Wet 36 5.8 38 4.5 39 3.8 38 3.8 38 4.0 36 3.1 36 2.9

5 58 Evap 112 30 133 25 155 22 163 11 174 8.0 153 11 159 12

(114,115) Wet 38 6.3 41 4.8 40 3.6 42 4.2 40 4.2 41 3.4 40 3.2

6 63 Evap 92 27 98 19 102 14 102 9.3 105 5.4 93 11 89 9.4

(116,117) Wet 47 10 48 8.0 49 7.2 47 9.1 47 9.1 46 6.7 46 6.3

7 65 Evap 102 30 128 28 135 22 145 10 151 3.1 142 5.6 136 9.3

(101,102) Wet 38 4.9 39 3.8 39 3.6 38 4.0 38 4.4 37 3.1 37 2.9

8 67 Evap 91 28 97 21 99 17 100 14 101 12 94 13 95 12

(104,105) Wet 42 9.8 43 7.0 45 6.4 43 7.7 42 8.3 38 5.4 42 5.6

9 69 Evap 117 30 164 18 163 19 170 6.1 172 3.8 164 8.7 169 4.9

(92,93) Wet 38 3.1 40 2.6 40 1.9 40 1.9 40 1.9 39 1.7 39 1.7

10 71 Evap 117 30 154 13 151 24 154 6.6 160 5.4 147 11 135 4.5

(95,96) Wet 39 3.5 41 2.6 39 1.9 41 2.5 40 2.0 40 2.0 40 2.0

11 73 Evap 116 30 141 27 150 22 157 9.2 168 5.3 157 7.7 154 8.4

(98,99) Wet 40 4.0 41 2.8 42 2.6 42 2.5 40 2.5 41 2.0 41 2.0

12 75 Evap 91 30 122 28 125 23 132 11 132 11 116 15 102 15

(124,125) Wet 45 15 47 10 47 9.0 46 12 45 13 44 9.4 45 8.9

13 87 Evap ...........................

(110) Wet 37 4.6 39 3.5 39 2.6 39 2.5 39 2.5 39 2.2 38 2.0

14 89 Evap 114 30 175 22 168 24 190 8.9 192 5.0 167 12 168 6.4

(119) Wet ............................

15 91 Evap 113 30 160 16 151 28 161 11 166 5.5 153 14 160 7.8

3.7

(122,123) Wet 40 6.3 43 4.5 42 3.9 43 4.7 43 3.4 43 3.9 43

NASA/TM--2001-210907 8
.5 y

o.Ts"@(

1.0 (,_,,___

1.5"

Figure 1: Cross-section of the airfoil leading

edge and heated zones

(_) 0.0005"

O.O08 / _)

0,0"22 ,/-.o_ ..............

0.035 ,,1.-=._

/

/

0,135 ,/-¢m

VVVVVVVVVVVVVV_

PVVVVVVVVVVVVVV

VVVVVVVVVVVVVV_

vVVVVVVVVVVVVVV

_v_vv_vvvvvvv'

f

Figure 2: Leading edge material composition

(see Table I)

50

30

.w

O"

10

-10

40 60 80 100 120

4- evap

Surface Temperature (deg. F)

_- cony 4- sens _ kinetic-m- total

Figure 3: Energy balance terms at _ot = 20 °F

5O

-10

40 60

30

O10

-e- evap

80 100 120

Surface Temperature (deg. F)

,-_ cony 4- sens _- kinetic -e- total

Figure 4: Energy balance terms at _ot = -22 °F

9O

70

50

"5

•r 30

g

10

-10

40

J

60 80 1O0 120

Surface Temperature (deg. F)

4- evap _,=--cony -m- sens _ kinetic

Figure 5: Energy balance terms at _ot = 20 °F in

percentage of total heat required

9O

7O

g 5o

_. ao

10

-10

40

7L 7

mL ................................. ].........

6O 80 100 120

Surface Temperature (deg. F)

evap -,=- cony -.m-sens _ kinetic

Figure 6: Energy balance terms at _or = -22 °Fin

percentage of total heat required

NAS A/TM--2001-210907 9
0.7

o_"0.6

0.5

]__ 0.4

,,=,

0.3

0.2

80.1

0

-5

..... ,..... ....... _............... _............................ ................

I i......... ..................... i i

-4 -3 -2 -1 0 1 2 3 4 5

S (in)

m 100mph _ 200mph

Figure 7: Predicted collection efficiencies Figure 10: Typical runback in a running-wet antiicing case at cold ambient temperatures

40

_ ao

N 20

¢-

10

50

....... i......

o ...... |!!!

-4 -3 -2 -1 0 1 2

Wrap Distance S (in)

l-I- Upper 4- Lower _ Average _ LEWlCE I

-T--i

3 4

Figure 8: Experimental and LEWICE predictions

of h at V= 1OOmph (Run#22, TTMS#51)

6O

5O

40

_- 30

g2o

t--

10

0

-4 -3 -2 -1 0 1 2 3

Wrap Distance S (in)

I-ii.-Upper -i-Lower .._-Average-I--LEWIOE I

__LLJ._LL_L_ ._i____L_i....J ___ _

_L_L....I..._....I........[.......I........ I.__,\.. J_ vLi_____. ,_- I I I i I I l/\_.\I

_,L_.J._L.I.._..i_-.ii_ ,_

I I..I......I_.L_L_L ............. i_

! II ! I I i

Figure 9: Experimental and LEWICE predictions

of h at V=2OOmph (Run#13, 7-1"MS#42)

Figure 1I: Typical runback in a running-wet antiicing case at warm ambient

tempera tures

6o

_.o I

20

'

10

o I i it °"

-6-5-4-3-2-1 0 1 2 3 4 5 6

Wrap Distance S, in

ANTtCE • E_p _ m

Figure 12: Experimental and ANTICEpredictions

for Run#22B (TTMS#71, running-wet)

0.7

1o.4

E

-0.3 -_

E

.0.2

NASA/TM--2001-210907 10
lOO

80

LL

40 T

20 ¢ Z

-4 -3 -2 -1 0 1 2 3

Wrap Distance S (in)

Figure 13: Experimental film coefficient for

Run#22B (TTMS#71, running-wet)

"8-

Upper

Lower

Averag_

LEWICE

Dry

r1

4

-6-5-4-3-2-1 0 1 2 3 4 5 6

Wrap Distance S, in

ANTICE • Exp _ m

Figure 16: Experimental and ANTICE predictions

for Run#67B (TTMS# 105, running-wet)

oo-F _, oo ' l

5O 0.8 5O

1.5

30 _ 30 1

Y_o _o.4 _ _"_o I o._"E

10 _0.2 10

0 ........ 0 0 0

-6-5-4-3-2-1 0 1 2 3 4

Wrap Distance S, in

ANTICE • Exp _ m

5 6

Figure 14: Experimental and ANTICE predictions

for Run#48B (TTMS#76, running-wet)

-7-6-5-4-3-2-1 0 1 2 3 4 5 6 7

Wrap Distance S, in

ANTICE • Exp _ m

Figure 17: Experimenta/ and ANTICE predictions

for Run#87 (7-1"MS#I 10, running-wet)

6O

_. 40

d_

30 _r

v !,

t./)

1-20 __,t

10 _,

It

0 :j:I

-6-5-4-3-2-1 0 1 2 3

Wrap Distance S, in

_,_] I I'

4-"f l-i_ .o.o_ il _lllil_, .0.4 N_-

i W l tti-.o.__

!.i..i,T,l.l.I, .M.,.

456

--.e,.- AN'rICE • Exp _ m

Figure 15: Experimental and ANTICE predictions

for Run#53B ('I-I'MS#81, running- wet)

180

160 ...................................................................................... _.............Upper {

140 _'

= _ i ,owor _----'120 ..............

_100 _,:_ Averag{

x: 60

40 _ _ Dry

2O

0 4 ! I ! I ! I I

-z -3 -2 -1 0 1 2 3 4

Wrap Distance S (in)

Figure 18: Experimental film coefficient for

Run#22A (TTMS#70, evaporative)

NASA/TM--200I-210907 11
180 0.2

140 I I _ _II_I

_-- 120 ._.

100 _J#_ l __.

_w -0,1 E

80 ==

_- 60 ,i_ l E

40 t i _ i "0.05

20 .Att !1 i'__

o .},1,1. ,_t_l+l_.!,.t.2 o

-6-5-4-3-2-1 0 1 2 3 4 5 6

Wrap Distance S, in

ANTICE • Exp _ m

Figure 19: Experimental and ANTICE predictions

for Run#22A (TTMS#70, evaporative)

120_t I + ' I

8O

_oo ' i

,o lilt i._t

t-- f/ t

° i,_.+++f, i.

-20

-6 -5 -4 -3 -2 -1 0 234

Wrap Distance S, in

0

5 6

0.8

0.6

t'-

o.4

v

E

0.2

ANTICE • Exp _ m

Figure 21: Experimental and ANTICE predictions

for Run#67A (TTMS#104, evaporative)

200 __.5

6_ 150 __.5

1oo

"--.-" _11.5 _

I-- E

50

-6-5-4-3-2-1 0 1 2 3 4 5 6

Wrap Distance S, in

ANTICE • Exp "-I1-" m

Figure 20: Experimental and ANTICE predictions

for Run#35A CTTMS#63, evaporative)

NASA/TM--2001-210907 12

REPORT DOCUMENTATION PAGE FormApprovea

OMB No. 0704-0188

Publtc reputing burden for this collection of information Is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources,

gathering and maintaining the data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of this

collection of information, including suggestions for reducing this burden, to Washington Headquarters Services, Directorate for Information Operations and Reports, 1215 Jefferson

Davis Highway, Suite 1204, Arlington, VA 22202-4302, and to the Office of Management and Budget, Paperwork Reduction Project (0704-0188), Washington, DC 20503.

1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED

May 2001

4. TITLE AND SUBTITLE

Validation of NASA Thermal ice Protection Computer Codes

Part 3: The Validation of Antice

6. AUTHOR(S)

Kamel M. Al-Khalil, Charles Horvath, Dean R. Miller, and William B. Wright

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

John H. Glenn Research Center at Lewis Field

Cleveland, Ohio 44135-3191

9. SPONSORING/MONITORiNG AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546-0001

Technical Memorandum

5. FUNDING NUMBERS

WU-711-21-23-00

8. PERFORMING ORGANIZATION

REPORTNUMBER

E-12786

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA TM--2001-210907

AIAA-97--005I

11. SUPPLEMENTARY NOTES

Prepared for the 35th Aerospace Sciences Meeting and Exhibit sponsored by the American Institute of Aeronautics and

Astronautics, Reno, Nevada, January 6-7, 1997. Kamel M. A1-Khalil and Charles Horvath, Cox & Company, Inc.,

200 Varick Street, New York, New York 10014; Dean R. Miller, NASA Glenn Research Center; and William B. Wright,

NYMA, Inc., 2001 Aerospace Parkway, Brook Park, Ohio 44142. Responsible person, Dean R. Miller, organization

code 5840, 216-433-5349.

12a. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified - Unlimited

Subject Category: 03 Distribution: Nonstandard

Available electronically at lgtp:J/gltrs, grc.nasa, gov/GLTRS

This publication is available from the NASA Center for AeroSpace Information, 301--621-0390.

12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum 200 words)

An experimental program was generated by the Icing Technology Branch at NASA Glenn Research Center to validate two

ice protection simulation codes: LEWICE/Thermal for transient electrothermal de-icing and anti-icing simulations, and

ANTICE for steady state hot gas and electrothermal anti-icing simulations. An electrothermal ice protection system was

designed and constructed integral to a 36 inch chord NACA0012 airfoil. The model was fully instrumented with thermocouples, RTD's, and heat flux gages. Tests were conducted at several icing environmental conditions during a two week

period at the NASA Glenn Icing Research Tunnel. Experimental results of running-wet and evaporative cases were

compared to the ANTICE computer code predictions and are presented in this paper.

14. SUBJECT TERMS

Aircraft ice protection; Thermal ice protection

17. SECURITY CLASSIFICATION

OF REPORT

Unclassified

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Unclassified

NSN 7540-01-280-5500

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Unclassified

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