Harold E. Addy, Jr.

Glenn Research Center, Cleveland, Ohio

Myron Oleskiw

National Research Council Canada, Ottawa, Ontario, Canada 

Andy P. Broeren

Glenn Research Center, Cleveland, Ohio

David Orchard

National Research Council Canada, Ottawa, Ontario, Canada 

A Study of the Effects of Altitude on 

Thermal Ice Protection System Performance

NASA/TM—2013-216559

September 2013

AIAA–2013–2934
	
	
	








	



	

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Harold E. Addy, Jr.

Glenn Research Center, Cleveland, Ohio

Myron Oleskiw

National Research Council Canada, Ottawa, Ontario, Canada 

Andy P. Broeren

Glenn Research Center, Cleveland, Ohio

David Orchard

National Research Council Canada, Ottawa, Ontario, Canada 

A Study of the Effects of Altitude on 

Thermal Ice Protection System Performance

NASA/TM—2013-216559

September 2013

AIAA–2013–2934


	



	


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NASA/TM—2013-216559 1 

A Study of the Effects of Altitude on Thermal Ice 

Protection System Performance 

Harold E. Addy, Jr. 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Myron Oleskiw 

National Research Council Canada 

Ottawa, Ontario, Canada K1A 0R6 

Andy P. Broeren 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

David Orchard 

National Research Council Canada 

Ottawa, Ontario, Canada K1A 0R6 

Abstract 

A study of the effects of altitude on an aircraft thermal ice protection system (IPS) performance was 

conducted. The study included tests of an airfoil model, with a heated-air IPS installed, in an altitude icing 

wind tunnel at altitude and at ground level conditions. An altitude scaling method was used to simulate 

the altitude icing conditions at ground level. The study allowed thermal and mass transfer phenomena to 

be examined. Three different flight scenarios were tested: Descent, Cold Hold, and Warm Hold. The test 

results showed that heat transfer was accurately simulated using the scaling method. Results of the mass 

transfer, in terms of amount of ice accreted, were less conclusive, but indicated that mass transfer was not 

as well matched. 

Nomenclature 

c model chord 

d twice the model leading edge radius 

qdot power density 

r recovery factor 

s surface distance 

IPS ice protection system 

K inertia parameter 

K0 modified inertia parameter 

LWC liquid water content 

M Mach number 

MVD median volumetric diameter 

Mw water loading 

Pr Prandtl number 

Re Reynolds number 

ReReynolds number based on droplet diameter 

Tr recovery temperature 
NASA/TM—2013-216559 2 

Tst static temperature 

Vtas true air speed 

0 collection efficiency at stagnation 

droplet median volumetric diameter 

 ratio of specific heats for air 

/Stokes droplet range parameter 

air air viscosity 

air air density 

w water density 

1.0 Background 

Thermal IPSs use heat energy to prevent the buildup of potentially dangerous ice accretions on 

aircraft when flying in icing conditions. That heat energy must be supplied by the aircraft. As new aircraft 

are being designed for higher efficiencies, all energy-using systems on the aircraft are being re-examined 

with an eye toward reducing their energy needs, including thermal IPSs. A thermal anti-icing IPS is 

designed to evaporate all water impinging on critical exposed surfaces of an aircraft during flight in even 

the most extreme conditions. This mode of operation is, however, quite energy-intensive. One way in 

which energy requirements can be greatly reduced is if, under the more severe icing conditions, a thermal 

IPS is designed to warm the impinging water only enough to allow it to run completely off the aircraft or 

back to a noncritical area before freezing. Such operation, however, requires that energy requirements be 

more accurately calculated and evaluated. These energy requirements must account for the many aspects 

of the aircraft icing process including air temperature, air speed, cloud water content, as well as aircraft 

altitude. While the effects of altitude, in particular reduced air density and pressure, have always been 

taken into account in the design of thermal IPSs, a better understanding of these effects is needed so as to 

enable more exact design, testing, and evaluation of these systems. 

Current methods to calculate the heat energy and mass transfer involved with thermal IPS operation 

are based on empirical relations. Heat transfer coefficients are typically derived from experiments 

designed to measure heat transfer from flat plates or cylinders. By similarity, the mass transfer coefficient 

is then assumed to be a function of the heat transfer coefficient. While these approximations have been 

adequate in the past, a more rigorous validation is needed if accuracy is to be increased, particularly when 

scaling is applied to account for the effects of altitude. 

For these reasons, the National Aeronautics and Space Administration (NASA) and the National 

Research Council Canada (NRCC) have undertaken an investigation of thermal IPS operation. This study 

is the first conducted specifically to better understand the effects of altitude on the heat and mass transfer 

processes occurring during operation of a thermal IPS in flight icing conditions. NRCC provided test time 

in the Altitude Icing Wind Tunnel (AIWT) in Ottawa, Canada. This facility provided aircraft inflight 

icing conditions over a range of air speeds, temperatures, and pressure altitudes as well as inflight icing 

cloud liquid water contents and droplet size spectra. NASA provided a NACA airfoil model equipped 

with a piccolo tube heated-air IPS. The model was instrumented such that heated-air flow rates and 

temperatures could be measured and inner surface temperatures of the model’s leading edge could be 

monitored. 

1.1 Prior Research 

Although no studies directly comparing the performance of a thermal IPS operating at altitude with its 

performance at ground level conditions have been reported, various aircraft icing research efforts have 

been conducted in the past that are of interest to the current study. These include the development of 

scaling methods for icing of unprotected surfaces, heat transfer studies, and ground level tests of thermal 

IPSs. 
NASA/TM—2013-216559 3 

Aircraft icing scaling methods for unprotected surfaces have been in development for a number of 

years. These methods include both geometric scaling for icing tests of subscale models and icing 

condition scaling for icing wind tunnel tests where a particular test point or points do not fall within the 

icing envelope that can be produced by the tunnel. Anderson (Ref. 1), Kind and Oleskiw (Ref. 2), and 

Bartlett (Ref. 3) provide descriptions and evaluations of these methods. While the addition of a relatively 

large source of heat by a thermal IPS to the icing processes is a major difference, these methods 

developed for unprotected surfaces can be used as a basis for similar research directed toward surfaces 

with a thermal IPS. The scaling method employed in this study used select parameters from these 

methods and are noted in the background section below. 

Heat transfer is an important aspect of aircraft icing and has been the focus of a number of studies. In 

particular, air turbulence level and the transition from laminar to turbulent flow dramatically affect icing, 

particularly at near-freezing temperatures. This is of particular concern when testing in icing wind tunnels 

where turbulence levels are typically well above those observed in flight. Research conducted by Newton, 

et al. (Ref. 4), provided important information on heat transfer in flight. That research was further 

extended to an icing tunnel by Poinsatte, et al. (Ref. 5) Important conclusions from this research were 

that, while the turbulence levels measured in the icing tunnel were 5 to 7 or more times as high as that in 

flight, there was good agreement in heat transfer between flight and tunnel tests at the lower Reynolds 

number conditions and only a slight increase in heat transfer at the higher Reynolds number conditions. 

Moreover, the increase in icing tunnel turbulence due to the addition of spray bar air did not change heat 

transfer over the tunnel operating without spray bar air. A later study by Bragg, et al. (Ref. 6) concluded 

that the addition of water to the spray did not appreciably change tunnel turbulence over that with the 

spray bar air only. These studies indicate that icing tunnel turbulence should not be a major factor in the 

current study. 

Miller, Wright, and Al-Khalil (Refs. 7 to 9) conducted tests in the NASA Icing Research Tunnel 

using an electro-thermal IPS directed toward validating NASA thermal IPS simulation codes. This effort 

found that the codes were fairly accurate in predicting airfoil surface temperatures when operating an 

airfoil model with IPS in icing conditions. It was noted that water evaporation and its effect on surface 

heat and mass transfer is important, as it is in this study, and warranted further research. 

1.2 Facility Description 

The AIWT is a specialized closed-loop low to moderate speed wind tunnel used to simulate in-flight 

atmospheric icing conditions (see plan view in Fig. 1). It is employed for internal research, for 

collaboration with other research organizations, and to assist commercial clients with: 

 Development, testing or calibration of aircraft or cloud physics instrumentation; 

 Development and testing of de- and anti-icing systems; 

 Evaluation of ice accretion on non-protected aircraft components; 

 Validation of numerical ice accretion codes; 

 Studies on basic physical processes leading to ice accretion. 

The tunnel’s standard test section is 22.5 in. square and 6 ft long. The airspeed in this test section can 

vary from about 10 to 195 kn. Two inserts are available to reduce the height of the test section and thus 

increase the airspeed. The first insert reduces the height to 12 in. and leads to a maximum speed of 

320 kn. The second insert further reduces the height to 10 in. with a corresponding maximum airspeed of 

365 kn. Access panels in the tunnel walls, floor and ceiling provide rapid access to test articles as well as 

flexibility in their mounting in the test section. Plexiglas windows are commonly installed in the test 

section to enable photographic or video recording of ice formation and growth. Airspeed in the tunnel is 

computer-controlled using a variable frequency drive which provides power to the fan’s 600 hp motor. 

Test section flow uniformity and a relatively low turbulence level are enhanced through the use of a 

honeycomb structure and screen at the entry of the settling chamber. Figure 2 shows a view of the AIWT 

test section with the operator’s console. 
NASA/TM—2013-216559 4 
NASA/TM—2013-216559 5 

A heat exchanger located upstream of the tunnel’s settling chamber permits rapid changes of air 

temperature within the test section. The use of a three-way valve to control the flow of chilled 

trichloroethylene through the heat exchanger permits a high level of temporal stability of air temperature 

in the tunnel. The closed cell insulation surrounding the tunnel shell assists in the ability to obtain static 

air temperatures as low as –30 F (–34.4 C). A thick steel tunnel shell combined with the operation of 

vacuum pumps permits partial evacuation of the air from the tunnel to simulate flight at altitudes as high 

as 30,000 ft. 

Six spray bars are located at the entry to the settling chamber, just downstream of the turbulencereduction screens. Up to 5 of 7 spray nozzles in each spray bar may be selected at any one time in various 

patterns to ensure adequate spray coverage across the test section. By varying the flow rate of distilled 

water to the external-mix spray nozzles, the liquid water content (LWC) in the test section may be varied 

between 0.1 and 3.0 g/m3

. Controlling the spray air pressure permits the median volumetric diameter 

(MVD) of the spray droplet size distribution to vary between 10 and 100 μm. Additional details regarding 

the tunnel are provided in Oleskiw, et al. (Ref. 10). 

1.3 Model Description

The airfoil model used for this study was an 18-in. chord NACA-0018. A symmetric airfoil was 

selected for this fundamental study primarily because it reduces the number of complicating factors in the 

test. A NACA-0012 airfoil was originally considered, largely because it has been extensively and 

successfully used in many previous fundamental studies of aircraft icing physics. The size of the AIWT 

test section, however, limits the size of the model to about 18 in. in chord length. For a NACA-0012, an 

18-in. chord results in a leading edge that is too narrow to allow installation of both a thermal IPS and 

measurement instruments. A hybrid airfoil was considered which would have incorporated a 36-in. chord 

NACA-0012 leading edge and truncated aft body to result in airfoil with an 18-in. chord. Such a hybrid 

airfoil, however, results in aerodynamics over the aft portion which might affect ice formed by water 

running aft along the airfoil surface, known as runback ice. This was undesirable for this study, so the 

hybrid design was not used. Instead, a NACA-0018 airfoil was chosen. It is a symmetric airfoil with a 

wider leading edge, allowing more space for installation of a thermal IPS and measurement instruments. 

Its leading edge also closely matches that of a 36-in. chord NACA-0012, as shown in Figure 3. 

The 2D, straight wing model spanned the 22.5 in. wide AIWT from wall to wall. It was constructed of 

a 0.063 in. thick 7075-T6 aluminum skin on a 0.625 in. thick, 7075 aluminum rib and spar frame. There 

were two ribs, one at each end, and two spars, near airfoil maximum thickness, as shown in Figure 4. The 

trailing edge was also machined from 7075 aluminum. 

A heated-air piccolo tube IPS was designed for the model. The primary objective of the design was to 

provide a 2D heating system with minimal span-wise variation in the temperature and heat transfer 

profiles during testing. This would simplify modeling and analysis of the measurements. For these 

reasons, the forward edge of the full-span piccolo tube was located 1.25 in. from the inside of the leading 

edge and had a single, straight row of 42, 0.032-in. diameter holes spaced 31/64 in. apart. A steel diffuser 

was designed to direct the heated airflow to and around the inside of the leading edge. Four aluminum 

ribs were used to maintain the diffuser shape. A cross section of the leading edge of the airfoil showing 

the IPS design is shown in Figure 5. 
NASA/TM—2013-216559 6 
NASA/TM—2013-216559 7 

Heated air was supplied to one end of the piccolo tube where it was directed to the inner surface of the 

leading edge via small holes. The heated air was then directed along the inner surface of the leading edge 

by a symmetric diffuser. The heated air exited the model through exhaust pipes mounted to a spar at the 

aft of the leading edge. The heated exhaust air flowing from the upper surface was kept separate from that 

flowing from the lower surface by a horizontal wall extending from the aft of the piccolo tube to the spar. 

Each of these two exhaust flows were measured by a Coriolis flow meter located in the respective exhaust 

pipe. Manual valves situated downstream of the flow meters were adjusted to ensure proper flow rates 

maintained equal heating to both the upper and lower surfaces of the leading edge. 

Thermocouples were used to measure model surface and heated air temperatures. A total of 25 thin 

film (0.0005 in. thick), T-type thermocouples were installed on the inner surface of the leading edge using 

a thin layer (approximately 0.002 in. thick) of thermally-conductive epoxy (M-Bond). The layout of the 

thermocouples on the leading edge is shown in Figure 6. 

Three T-type thermocouples were used to monitor the temperature of the heated air flowing into the 

model. One was at the entrance to the piccolo tube while two were spaced axially along the center of the 

tube using a rod device that was inserted into the end of the piccolo tube. Heated air exhaust temperatures 

were measured using six T-type thermocouples mounted through the front spar as shown in Figure 7. 
NASA/TM—2013-216559 8 

2.0 Approach

2.1 Test Plan 

Icing conditions were selected to approximate three different aircraft icing scenarios at altitude: an 

aircraft hold at relatively low icing air temperature (Cold Hold), a hold at relatively warm icing 

temperature (Warm Hold), and a descent at intermediate icing temperature (Descent). These conditions 

are given in Table 1 and will be referred to as reference conditions throughout the remainder of this 

report. IPS heated air flow rates and temperatures were set to provide a ‘running wet’ mode of operation 

for the ice protection. In this mode, all the water impinging on the leading edge of the model is not 

evaporated. Some is allowed to flow aft before freezing in an area that is presumably less critical to the 

model’s aerodynamics. For the purposes of this test, however, aerodynamic criticality was not a primary 

consideration. Instead, the heat and mass transfer phenomena involved in the accretion of ice in 

conjunction with a heat source (a thermal IPS), is the focus of this study. 

The test plan also included runs without the IPS operating. These runs provided data in terms of 

unprotected ice accretion shapes and masses that could be used for post-test analysis and comparison 

purposes. 

TABLE 1.—REFERENCE CONDITIONS 

Flight phase Altitude, 

ft 

Vtas, 

kn 

AOA, 

deg 

Tstatic, 

F 

Ttotal, 

F 

LWC, 

g/m3

MVD,

m 

Descent 10,000 180 0 –4.0 3.7 0.15 20 

Cold Hold 15,000 180 0 –22.0 –14.3 0.15 20 

Warm Hold 15,000 180 0 16.5 24.2 0.50 20 

2.2 Scaling Method 

In addition to the three aircraft icing scenarios at altitude, three corresponding icing conditions at 

ground level were selected for the test. Because the heat and mass transfer associated with the icing 

process in the presence of a thermal IPS are known to be affected by air density and pressure, a method to 

scale for these effects was employed. This method matched the following four parameters at altitude and 

ground level for each of the three scenarios in the test: 

1. External Reynolds number 

2. Water loading 

3. Modified droplet inertia parameter 

4. Recovery temperature 
NASA/TM—2013-216559 9 

Empirical analysis has most often found that the heat transfer coefficient is typically a function of the 

external Reynolds number. Since, by similarity, the mass transfer coefficient is a function of the heat 

transfer coefficient; both coefficients are dependent on Reynolds number. It follows then, that to obtain 

similar results at altitude and ground level, the external Reynolds number must be the same for both 

reference and scale cases. For this study, the following expression for Reynolds number was used (Ref. 1): 

air

air tas Re



 	 	 
 V d (1) 

The amount of water impinging on the leading edge of a model in icing conditions also has a major 

effect on the ice that forms. Therefore, it is important to maintain the amount of water impinging on the 

model at the same level for both the altitude and ground level cases. The following expression for water 

loading was used (Ref. 1): 

LWC tas 0 Mw 
 	V 	 (2) 

The area of coverage of the impinging water should also be the same between altitude and ground 

level. This was achieved by matching the modified droplet inertia parameter (Ref. 1): 






 



  

 
 

8

1

8

1

Stokes

K0 K (3) 

where 

air

tas

2

18 

 	 	 
 dV K w (4) 

air 

  


 



0.8388 0.001483Re 0.1847 Re

1

Stokes

 (5) 

Air temperature also plays a major role in ice formation. Because matching external Reynolds number 

between altitude and ground level conditions results in a very different airspeed for the two, it then 

follows that both total air temperature and static air temperature cannot be matched. For unprotected 

surfaces, total air temperature is a driving factor, with static air temperature playing a lesser role. In a 

running wet situation, where the ice forms further aft on the airfoil, static air temperature may play a 

greater role. Therefore, for this test, it was decided to match recovery temperature, which is between total 

and static air temperature, but is closer to total. 
















 




 



   
  2

2

1 Tr Tst 1 r M (6) 

where 

r 
 Pr (7) 

The scaled conditions for each of the three flight scenarios along with the corresponding conditions at 

altitude (reference) are given in Table 2 and will be referred to as scale conditions throughout the 

remainder of this report. 
NASA/TM—2013-216559 10 

TABLE 2.—REFERENCE AND CORRESPONDING SCALE CONDITIONS 

Flight phase Alt., 

ft 

V, 

kn 

AOA, 

deg 

Tstatic, 

F 

Ttotal, 

F 

LWC, 

g/m3

MVD, 

m 

Red 

106

Mw,

g/m2-s 

K0 Tr, 

F 

Descent(ref) 10,000 180 0 –4.0 3.7 0.15 20.0 1.64 8.81 1.41 2.5 

(sc) 130 0 –0.9 3.1 0.21 24.5 1.64 8.81 1.41 2.5 

Cld Hld(ref) 15,000 180 0 –22.0 –14.3 0.15 20.0 1.40 9.02 1.52 –15.5 

(sc) 106 0 –17.8 –15.1 0.25 27.8 1.40 9.02 1.52 –15.5 

Wm Hld(ref) 15,000 180 0 16.5 24.2 0.50 20.0 1.29 30.0 1.50 23.0 

(sc) 105 0 20.8 23.4 0.85 27.8 1.29 30.0 1.50 23.0 

The reason for this approach was twofold: a) it permitted a close examination of the heat and mass 

transfer as well as the associated assumptions involved and b) it will aid in the development of altitude 

scaling methods for icing tests where altitude simulation is not available. 

2.3 Test Procedure 

Estimates for the thermal IPS’s heated air temperature and flow rate were made using past test 

experience and LEWICE predictions. These values were then adjusted slightly during checkout runs at 

the reference conditions to obtain runback ice shapes that were judged to be desirable for both recording 

and computational modeling purposes. The desired ice shape was a runback ridge of ice running 

continuously across the span of the leading edge of the airfoil. This ice shape can be predicted by 2D ice 

accretion codes such as LEWICE2D and can be traced on a template fit over the leading edge of the 

model in the tunnel. The IPS settings determined during checkout runs at the reference conditions were 

then also used for the scaled conditions. 

For each test run, the AIWT was first brought to the desired airspeed and air temperature. For the 

reference runs, AIWT pressure was set to obtain the desired altitude value. The model IPS operating 

conditions were also set at the predetermined settings. Once the tunnel and the model had reached steadystate operating conditions, the tunnel’s water spray system was activated, exposing the model to icing 

conditions for a predetermined period of time. 

At the conclusion of each test run, photographs were taken of the accreted ice and ice tracings were 

made to record the shape and location of the ice on the airfoil. Ice depth measurements were also made 

and then the ice was removed from the model and weighed. Three ice tracings were typically made for 

each ice accretion; one at the mid-span of the model and one each at 5.25 in. on either side of mid-span. 

This covered an area of relatively uniform LWC for most clouds used in these tests. Since the ice was cut 

at these locations, the ice mass values reported, unless otherwise noted, are for the ice accreted over the 

center 10.5 in. of the model. Results of the tests will be discussed in terms of these measurements as well 

as the model leading edge temperatures and heated air flow rates and temperatures.

Model and heated air thermocouples were calibrated before and after the tests and were found to be 

within 1 C. The heated air flow meters were recalibrated after the tests and were found to be accurate to 

1 percent. A root-mean-squared error analysis of the heated-air power density calculations yields less 

than two percent error. The error in ice depth and mass measurements can only be approximated as the 

techniques used were highly manual. Repeat measurements of ice accreted without the IPS operating 

indicated that they were likely to be within 10 percent. 

2.4 Pre-Test LEWICE Study 

Prior to running the tests at the AIWT, the LEWICE (Ref. 11) ice accretion code was used to predict 

the resulting ice shapes using the test conditions to help ensure that the approach and methods employed 

were sound. Figures 8 to 10 show the LEWICE results from the Descent, Cold Hold, and Warm Hold 

cases, respectively, for both the altitude and ground level conditions. The close agreement indicates that 

the scaling method is promising. It should be noted that results from this code using the thermal IPS 

simulation capability are highly dependent upon how the surface grid and boundary conditions are set up. 
NASA/TM—2013-216559 11 
NASA/TM—2013-216559 12 

3.0 Results 

3.1 Post Test Data Analysis 

During post-test calibration of the AIWT, it was found that the cloud operating settings produced 

droplet sizes that differed from those desired. The calibration found that the droplet sizes for the reference 

conditions in each case were smaller than intended and they were larger than intended for the scaled 

cases. This affected two of the four parameters that were to be matched in the scaling method. The actual 

values for modified inertia parameter, K0, and water loading, Mw, were no longer matched between their 

respective reference and scale values. Both the intended and actual test conditions and resulting scale 

parameters are given in Table 3. 

TABLE 3.—INTENDED AND ACTUAL (SHOWN IN BOLD) CONDITIONS 

Flight 

phase 

Alt., 

ft 

Vtas, 

kn 

AOA, 

deg 

Tstatic, 

F 

Ttotal, 

F 

LWC, 

g/m3

MVD, 

m 

Red

106

Mw

g/m2-s 

K0 Tr, 

F 

Descent(ref) 10,000 180 0 –4.0 3.7 0.15 20.0 1.64 8.81 1.41 2.5

10,000 180 0 –4.0 3.7 0.14 17.9 1.63 7.74 1.19 2.5

(sc) 130 0 –0.9 3.1 0.21 24.5 1.64 8.81 1.41 2.5

130 0 –0.9 3.1 0.22 28.1 1.59 9.98 1.75 2.4

Cld Hld(ref) 15,000 180 0 –22.0 –14.3 0.15 20.0 1.40 9.02 1.52 –15.5

15,000 180 0 –22.0 –14.3 0.15 16.2 1.39 8.03 1.10 –15.5

(sc) 106 0 –17.8 –15.1 0.25 27.8 1.40 9.02 1.52 –15.5

106 0 –17.8 –15.1 0.26 32.6 1.36 9.91 1.95 –15.5

Wm Hld(ref) 15,000 180 0 16.5 24.2 0.50 20.0 1.29 30.0 1.50 23.0 

15,000 180 0 16.3 24.0 0.48 16.3 1.29 25.6 1.09 22.8 

(sc) 105 0 20.8 23.4 0.85 27.8 1.29 30.0 1.50 23.0 

105 0 20.8 23.4 0.87 29.8 1.27 31.4 1.67 23.0 
NASA/TM—2013-216559 13 

The values for K0 were most affected with the scale values being greater than the corresponding 

reference values. For the Descent case, the K0 scale value was 42.8 percent greater than the reference 

value; for the Cold Hold case, it was 77.3 percent greater; and the Warm Hold case, it was 50.0 percent 

higher. The scale values for Mw were also greater than there corresponding reference values. In the 

Descent case, it was 28.9 percent higher; in the Cold Hold case, it was 23.4 percent higher; and in the 

Warm Hold case, the Mw scale value was 21.6 percent higher than the reference value. 

Of these two parameters, K0 and Mw, Mw is likely to have a greater influence on the results. K0 affects 

the droplet impingement limits and the collection efficiency, but Mw drives the rate of ice buildup or, in 

the instance of an IPS, the amount of power required to prevent or dramatically reduce ice buildup. 

The results of this study, therefore, are less clear than they would have been otherwise. Nonetheless, 

this test, being the first of this particular nature, is of interest and, moreover, conclusions can still be 

drawn, albeit with slightly less certainty. 

3.2 Descent Case 

Figures 11(a) and (b) show photographs of the ice accreted on the model for both the reference and 

scale conditions for the Descent case. Qualitatively, the ice shapes are similar. This similarity is reflected 

in the mid-span ice tracings shown in Figure 12. The tracings do, however, show that more ice was 

accreted at the scale conditions and that it froze slightly further aft on the model. The measured masses of 

ice and the height of the ice on both the upper and lower surfaces for both reference and scale conditions 

are given in Table 4. 
NASA/TM—2013-216559 14 

TABLE 4.—DESCENT CASE RESULTS 

Flight Phase Alt., 

ft 

Vtas, 

kn 

AOA, 

deg 

Tstatic, 

F 

Ttotal, 

F 

LWC, 

g/m3

MVD, 

m 

, 

s 

Up ice, 

in. 

Lo ice, 

in. 

Ice 

mass, 

g 

Descent(ref) 10,000 180 0 –4.0 3.7 0.14 17.9 600 0.14 0.18 29 

Descent(sc) 130 0 –0.9 3.1 0.22 28.1 600 0.18 0.22 54 

Figure 13 shows the model surface temperatures during the run both before exposing the model to the 

cloud (dry) and again during exposure to the cloud (wet). The close agreement between the reference and 

scale in the dry conditions indicate the degree to which matching Reynolds number matches heat transfer 

between the two. There is more of a difference between the surface temperatures in the wet conditions. 

Here, the confounding effects of differing Mw and K0 make it difficult to ascertain the cause. It is expected 

that the increased loading of supercooled water in the scale conditions would result in lower surface 

temperatures, but this is not the case. Instead, increased surface cooling was observed during the wet 

reference conditions. Increased evaporative cooling is expected at the reference conditions due to the 

lower tunnel air pressure. Mathematical models to calculate the amount of evaporative cooling in this 

situation, however, are approximate and have not been validated. 

The measured reference and scale IPS heated air inlet and outlet temperatures for both the dry and wet 

conditions for this case are shown in Figure 14(a). Similarly, the heat energy used by the IPS is shown in 

Figure 14(b). The close agreement between reference and scale values for these parameters even though 

Mw was 28.9 percent higher and K0 was 42.8 percent higher in the scale conditions shows that this IPS is 

not greatly affected by such changes in external icing conditions. 

During pre-test analysis, calculations indicated that the water evaporation rate would be slightly higher 

for the reference conditions than in the scale conditions. Further calculations showed that the evaporation 

rate for the scale conditions could be raised by slightly increasing the surface temperatures of the model 

by increasing the heated air temperature supplied to the model IPS. A test run was, therefore, made with 

the surface temperatures increased by 2.8 C (5.0 F). Results of this test run are shown in the photograph 

in Figure 15 and the ice tracings in Figure 16. The resulting scale ice accretion was reduced slightly in 

size and mass, but is still different from the reference ice accretion. The ice was also accreted further aft 

on the model than in the initial scale run. The desired scale-condition ice accretion is one that matches the 

reference ice accreted in size, shape, and location. As expected, the surface temperature profiles, shown in 

Figure 17, as well as the heated air inlet and outlet temperatures and heat energy plots, shown in Figure 

18, reflect the increased heated air supply temperature. 
NASA/TM—2013-216559 15 
NASA/TM—2013-216559 16 

3.3 Cold Hold Case 

For tunnel operation purposes, the scale conditions were run first for this Cold Hold case only. No 

checkout run was performed for the reference conditions, therefore, the reference ice accretion was not 

the result of adjusting the IPS settings to obtain a more desirable accretion. The case was not repeated due 

to time and priority considerations. 

Photographs of the ice accreted for the reference and scale conditions of the Cold Hold case are shown 

in Figure 19. The mid-span ice tracings are shown in Figure 20. Unfortunately, a relatively large piece of 

ice shed from the upper surface of the model during the run at reference conditions. The shed ice prevents 

a quantitative comparison of the upper surface ice thickness, and overall mass. The ice on the lower 

surface was fairly similar, however, and it is not unreasonable to assume the upper ice accretions would 

also have been similar had the piece of ice not been shed. It will be noted that the ice accretion formed 

more forward in these runs. The measured ice height and mass data for this case are given in Table 5. 
NASA/TM—2013-216559 17 

TABLE 5.—COLD HOLD CASE RESULTS 

Flight phase Alt., 

ft 

Vtas, 

kn 

AOA, 

deg 

Tstatic, 

F 

Ttotal, 

F 

LWC, 

g/m3

MVD, 

m 

, 

s 

Up ice, 

in. 

Lo ice, 

in. 

Ice 

mass, 

g 

Cld Hld(ref) 15,000 180 0 –22.0 –14.3 0.15 16.2 600 0.12 0.20 25 

Cld Hld(sc) 106 0 –17.8 –15.1 0.26 32.6 600 0.20 0.23 56 

Model surface temperatures for this case are shown in Figure 21. The dry model temperature profiles 

are nearly the same for the reference and scale icing conditions, again indicating similarity in heat 

transfer. For icing cloud on, wet surface conditions, the reference surface temperature profile, as in the 

Descent case, is lower than the profile for the scale conditions. Here, Mw is 23.4 percent higher and K0 is 

77.3 percent higher in the scale conditions. Figure 22 shows the heated air temperatures and heat energy 

used by the IPS. Once again, IPS operating parameters are not greatly affected by such deviations from 

the target icing condition parameters, Mw and K0. 
NASA/TM—2013-216559 18 

3.4 Warm Hold Case 

Warmer ambient air temperatures typically result in ice accretions of less structural strength while 

higher water loading conditions usually result in larger ice accretions. The combination of these two 

factors made the Warm Hold case the most challenging for which to obtain measurable runback ice. 

Several initial runs at the reference conditions were made during which large portions of ice were shed 

before reaching the end of the designated 10 min icing exposure period. The exposure time was then 

reduced to seven minutes and a run at reference conditions was completed without ice shedding. A 

photograph of the ice accreted at the reference conditions, along with the centerline ice tracing, are shown 

in Figures 23(a) and (b), respectively. When the scale conditions were run, however, the runback ice did 

not form a ridge. It instead froze in the form of rivulets further aft on the airfoil. Photographs of this ice 

are shown in Figure 24(a) and (b). This frozen rivulet ice did not lend itself to the tracing method used for 

the runback ridge ice. The tracing templates did not extend far enough aft in the chordwise direction to 

capture this ice. Moreover, frozen rivulet ice is oriented primarily in the chordwise direction, with 
NASA/TM—2013-216559 19 

relatively large spaces between the rivulets in the spanwise direction. Because the tracing templates were 

not designed to trace this type of ice, no tracings were made. The thickness of the frozen rivulets was, 

however, measured using a depth gage in the forward part of the rivulet. The ice was removed from the 

model and weighed, similar to the instances where ridge ice formed; however, all of the rivulet ice was 

removed and weighed, instead of weighing only the ice from the center section 10.5 in. section of the 

airfoil. The ice depth and mass data for this case are shown in Table 6. 

For this case, the post-test tunnel cloud calibration yielded an Mw that was 21.6 percent higher and a 

K0 that was 50.0 percent higher for the scale conditions. It appears likely that these deviations in cloud 

conditions resulted in dramatically different ice accretions. Although verification of this observation is 

warranted, it would illustrate the highly non-linear nature of icing at near-freezing temperatures. 

TABLE 6.—WARM HOLD CASE—RIDGE ICE RESULTS 

Flight phase Alt., 

ft 

Vtas, 

kn 

AOA, 

deg 

Tstatic, 

F 

Ttotal, 

F 

LWC, 

g/m3

MVD, 

m 

, 

s 

Up ice, 

in. 

Lo ice, 

in. 

Ice 

mass, 

g 

Wm Hld(ref) 15,000 180 0 16.5 24.2 0.48 16.6 420 0.48 0.51 53 

Wm Hld(sc) 105 0 20.8 23.4 0.87 29.9 420 ---- ---- ---- 
NASA/TM—2013-216559 20 

The airfoil surface temperature data are given in Figure 25. As in the other cases, surface temperature 

profiles are the same for the dry conditions and close for the wet conditions with the reference-run 

temperatures slightly lower (1 C or less). Heated air inlet and outlet temperatures as well as the heat 

energy input for the reference and scale conditions are nearly identical, as shown in Figure 26. 

The difficulties associated with successfully accreting and recording runback ridge ice accretions 

became a major obstacle to obtaining results for use in the study. Therefore, the test plan was changed to 

instead accrete and record runback rivulet ice accretions as the reference for the Warm Hold case. This 

was done by raising the heated air inlet temperature by 20 F (11.1 C). While this eliminated ice tracings 

as a means of recording and comparing test results, the height of the frozen rivulets could still be 

measured and the ice could still be removed and weighed for comparison purposes. 
NASA/TM—2013-216559 21 

Photographs of the resulting ice from the test runs at the Warm Hold conditions with runback rivulet 

ice accretions are shown in Figure 27. An examination of the frozen rivulets formed shows that, similar to 

the cases run with runback ridge ice, more ice accreted at the scale conditions than at the reference 

conditions. This increase in accreted ice is reflected in the measured ice heights and masses given in 

Table 7. During close observation of the formation of the frozen rivulets, beads of water were seen 

running aft along the length of the rivulets and then disappearing. It is suspected that these relatively large 

beads of water were being re-entrained into the airstream, rather than freezing and adding to the ice mass 

on the airfoil. 

Figure 28 shows the model surface temperatures for this case. They are again close between the 

reference and scale runs, indicating a close matching of heat transfer between the two. Figures 29(a) and 

(b) show the heated air data for this case. The close matching between reference and scale conditions also 

substantiate a matching of heat transfer as well as a reduced level of sensitivity of the IPS operation to 

deviations in cloud parameters. 
NASA/TM—2013-216559 22 

TABLE 7.—WARM HOLD CASE—FROZEN RIVULET ICE RESULTS 

Flight phase Alt., 

ft 

Vtas, 

kn 

AOA, 

deg 

Tstatic, 

F 

Ttotal, 

F 

LWC, 

g/m3

MVD, 

m 

, 

s 

Up ice, 

in. 

Lo ice, 

in. 

Ice 

mass, 

g 

Wm Hld(ref) 15,000 180 0 16.5 24.2 0.48 16.3 420 0.08 0.09 15 

Wm Hld(sc) 105 0 20.8 23.4 0.87 29.8 420 0.13 0.12 83 

3.5 Discussion

For all cases, the close matching of the surface temperatures and heated air temperatures and energy 

use rates between the reference and scale conditions showed that the scaling method works well in 

matching the heat transfer between altitude and ground level conditions. The fact that there was 

consistently less ice accreted for the reference conditions than for the corresponding scale conditions, 

however, suggested that the scaling method does not match mass transfer between altitude and scale 

conditions. Confounding this observation is the post-test calibration finding that Mw was 21.6 to 

28.9 percent higher for the scale conditions in each flight-scenario case. In general, the mass of the 

accreted ice was roughly 80 to 500 percent higher for the scale conditions in each case. The additional ice 

accreted in the scale conditions cannot be due to the increase in Mw unless the ice accretion rate in each 

case was highly nonlinear, a distinct possibility. 

The one exception to mass transfer scaling mismatch is the Cold Hold case. Here ice which shed from 

the upper surface of the reference conditions precluded a direct comparison of the ice masses between 

reference and scale. Inspection of the ice tracings, however, indicated that the masses may have been 

similar had the ice not shed. The additional ice accreted due to the increased Mw in the scale conditions 

would still have placed the ice masses within the margin of error estimated for these measurements. It 

should further be noted that the ice accreted for this case nearly covered the entire leading edge, thereby 

reducing the area for evaporation as well as reducing the length of runback for the water. The case 

rightfully should have been rerun with a higher heated air energy input such that the IPS operated more 

effectively. As was mentioned earlier, tunnel test time and run priorities prevented a repeat run of this 

case. One should be cautioned against drawing conclusions from this single case. 

All of the Descent and Warm Hold cases with the IPS operating resulted in less ice accreted during 

runs with the reference conditions than during the runs with scale conditions. This was evident even 

though ice shedding was a problem for the Descent cases and the Warm Hold cases where runback ridge 

ice was formed. To further illustrate the ice accretion differences between reference and scale conditions, 

normalized icing rates for the Descent and Warm Hold conditions are plotted in Figures 30 and 31, 

respectively. Icing rates, in terms of grams of ice per unit area per unit time, were calculated from the ice 
NASA/TM—2013-216559 23 

mass measurements, the pertinent impingement area of the airfoil, and the icing exposure time. The 

normalized values were obtained by dividing the icing rates for the runs with the IPS operating by the 

rates for the corresponding unprotected runs. Figure 31 shows multiple Warm Hold runs for both 

reference and scale conditions because the frozen rivulet results proved to be more repeatable than the 

runback ridge ice results where shedding ice was a frequent occurrence. For both cases, the increase in 

accreted ice for the scale conditions was greater than the unintended increase in Mw. 

The difference between the reference and scale ice accretions is more pronounced for the Warm Hold 

case than for the Descent case. It is strongly suspected that a reason for at least part of difference between 

the Warm Hold and Descent cases is the observation of the runback water droplets on the frozen rivulets 

which appeared to be re-entrained into the airstream. This phenomenon was not readily noticeable during 

the Descent cases where only runback ridge ice was formed. 

The unintended increase in Mw and K0 for the scale conditions tested in this study added uncertainty to 

the results. Those results, nonetheless, strongly suggested that, in order to faithfully simulate the amount 

of runback ice formed where a running wet IPS is being tested at ground level, another scaling parameter 

may well need to be considered or an appropriate correction method be employed. Methods used to 

calculate water evaporation rates should also be examined more closely. The evaporation rates used 

during prediction of ice accretion on unprotected surfaces are relatively low in comparison to the 

incoming water, so errors in the evaporation rate calculation are not critical. This is not the case for a 

thermal IPS that is operating effectively. Here a relatively large portion of the incoming water is 

evaporating, and thus accurate calculation is more important when determining the amount of runback ice 

that accumulates. 

In cases of aircraft icing where flow of water is known to occur, such as during the accretion of glaze 

ice on an unprotected surface, the Weber number can be an important factor in scaling icing conditions to 

obtain ice shape similarity. The flow of water is obviously present in the instance of frozen rivulet icing. 

Therefore, the inclusion of the Weber number should be considered in future research. 
NASA/TM—2013-216559 24 

4.0 Closing Remarks

This study was the first documented effort directed specifically toward better understanding the effects 

of altitude on the heat and mass transfer processes occurring during operation of a thermal IPS. The tests 

described should be considered an initial step toward providing technology that is essential to the 

enhancement of aircraft thermal IPSs and, thereby, aviation safety. Further steps in the study of the effects 

of altitude on IPS performance should include testing in an altitude icing wind tunnel to: 

1. Verify initial findings, 

2. More firmly establish the existence of water droplet re-entrainment, and investigate methods to 

quantify it, and 

3. Establish methods to determine water evaporation rates. 

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Testing at Reduced Scale,” ICAS 2002 Congress, 2002. 

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“Validation of NASA Thermal Ice Protection Computer Codes Part 1-Program Overview,” AIAA97-0049, NASA TM–107397, 1997. 

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Computer Codes: Part 2 – The Validation of LEWICE/Thermal,” AIAA-97-0050, 1997. 

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Altitude Icing Wind Tunnel,” AIAA-2001-0094, 2001. 

11. Wright, W., “User’s Manual for LEWICE Version 3.2,” NASA/CR—2008-214255, 2008.