NASA Contractor Report 175089

Wind Tunnel Tests of Rotor Blade

Sections With Replications of Ice

Formations Accreted in Hover

(NASA-CE-175089) HIN D TUNNE L TESTS OF'fiOTOB

BLADE SECTIONS KI3H REPLICATIONS OF ICE

FORMATIONS ACCBETED IN HOVEfi Final Report

{Ohio State Oci-v.) .31 p HC AG3/MF A01

CSCL ,01C G3/03

John D. Lee

The Ohio State University

Aeronautical and Astronautical Research Laboratory

Columbus, Ohio

and

Jack H. Berger and Timothy J. McDonald

FluiDyne Engineering Corporation

Columbus, Ohio

N86-22558

Unclas

059.17

March 1986

Prepared for

Lewis Research Center

Under Grant NAG 3-374

NASA

National Aeronautics and

Space Administration
TABLE OF CONTENTS

SUMMARY

NOMENCLATURE

INTRODUCTION

WIND TUNNEL MODELS

FACILITY DESCRIPTION

TEST PROGRAM

RESULTS AND DISCUSSION

APPENDIX: EVALUATION OF TUNNEL INTERFERENCE 9

REFERENCES 11
SUMMARY

Full scale reproductions of ice accretions molded

during the documentation of a hover test program were fabricated

by means of epoxy castings and used for a wind-tunnel test

program. Surface static pressure distributions were recorded

and used to evaluate lift and pitching moment increments while

drag was determined by wake surveys. Through the range of the

tests, corresponding to those conditions encountered in hover

and in flat pitch, integration of the pressure distributions

showed negligible changes in lift and in pitching moment, but

the drag was significantly increased.
NOMENCLATURE

C Chord Length (0.827 meter or 1.750 feet)

C Drag Coefficient

Cr Lift Coefficient LI

C Pitching Moment Coefficient (about quarter-chord)

C Pressure Coefficient

LWC Icing Cloud Liquid Water Content (gm./m )

M Mach Number

R Radius of Rotor (7.315. m., 24 ft.)

r Radial Distance Along Rotor

X Chordwise Distance from Leading Edge

Attack Angle
INTRODUCTION

A s par t o f th e join t NASA/Arm y HIF T progra m

(Helicopte r Icing Flight Test), ice was accreted on the main

rotor of a UH-1H helicopter during several hover flights at the

Canadian hover spray rig at Ottawa and documented upon landing.

The program and the documentation procedures are described in

reference s 1,2 and 3. Table I lists the pertinent information

for the five flights whic h wer e successfully documented. In

orde r to determine the aerodynamic characteristics of the blade

with the ice, a v/ind tunne l test program was planned and executed using cast replicas of the ice formations on full scale

models.

One of the methods for documenting the ice shapes, a

silicone-rubber molding technique, was so successful in capturing the fin e details of the ice formations that it provided the

basis for the wind tunnel program. Molds were obtained at up to

eight spanwise stations on the blade, depending on the extent of

the ice, wit h each mold bein g abou t 10 inche s in lengt h

(spanwise) . The last of the fiv e flights (fligh t E) wa s

selected as having the greatest potential for correlation betwee n th e measure d aircraf t performanc e an d performanc e

calculations based on section aerodynamics. For that flight ,

Figur e 1 shows the projections fro m the ends of the silicone

molds, from which the accreted ice-shapes could be reproduced in

some detail. The spanwise distribution of the ice is presented

in Figure 2.

Models wer e fabricate d b y repetitivel y castin g

replicas of the ice formations from the molds onto sections of

UH-1H helicopter blades which were then pressure tapped in the

front 35% (of chord) to acquire pressure distributions. The wind

tunne l test progra m wa s performe d b y FluiDyn e Engineerin g

Corporation (Minneapolis) in their 66-inch transonic tunnel [4].
This report summarizes the wind tunnel program and the

results obtained from it.

WIND TUNNEL MODELS

Four wind tunnel models were fabricated using 66 inch

sections of UH-1H main rotor blades supplied by the Army; for

referenc e purposes, the section was basicall y a NACA-001 2

profile . The ice formations were reproduced in epoxy castings

fro m the silicone molds directly onto the blade-model ; the

casting was performed repetitively using a single mold, starting

v/ith the first casting spanning the centerline and then alternatin g on eithe r side to for m a 66 inc h span quasi-two -

dimensional model. This procedur e was followed in orde r that

the best reproduction would be at the mid-span in case any mold

material was lost in removing it fro m each epoxy casting; the

detail of the reproductions indicated that this precaution was

unnecessary.

Pressur e taps wer e installed at the mid-span of the

model by drillin g throug h to the D-spar and inserting steel

hypodermi c tubin g throug h the hole and out the end of the Dspar. This procedure was possible for only the front 35% of the

models; however, some preliminar y win d tunne l tests in scale

indicated tha t the pressur e field distortio n due to the ice

accretion shape was confined to the front 35% or less, and the

full-scale data later confirmed this. Thre e taps wer e located

on or beside the main ice formation and thereafte r at 5% intervals throug h 35% . Althoug h dat a fro m suc h a se t wa s

insufficien t to defin e the details of the flowfield , it was

considered adequate for calculating changes in lift and pitching

movement.
Models were fabricated for 4 stations (r/R = 0.44,

0.60, 0.77 and 0.93) as being representative of the blade in

providing a minimum amount of data for evaluating the rotor

performance. A "clean" blade model provided baseline reference

data for the 3 inner stations as well as primary data for r/R =

0.93 since no ice was accreted that far out for flight E. The

tap layout for the four models is given in Figure 3.

For two stations (r/R = 0.44 and 0.60) an attempt was

made to simulate the ice formations with simple strips of wood

cut, in cross-section, in a trapezoidal shape approximating the

shape of the main ice formation. These were attached, in turn,

to the clean model. The wood strips were not tapped and so the

pressure distributions from those tests were incomplete in the

vicinity of the leading edge.

FACILITY DESCRIPTION

The full-scale rotor-blade models were tested in the

66 in. X 66 in. transonic wind tunnel of Fluidyne Engineering

Corporation, Minneapolis. This is a slotted-wall tunnel driven

by air ejectors; it has an atmospheric intake which can be

heated to avoid moisture condensation (see Figure 4) . A multiple tube wake-rake v/as used for the determination of drag.

Figure 5 shows the arrangement for mounting the models in the

wind tunnel and the location of the wake rake ( at 0.75 chord

length downstream from the trailing edge ). A photograph of a

model mounted in the tunnel is given in Figure 6. In a typical

test sequence, the test section Mach number was attained by

adjusting the ejector flow rate, the model pressures and the

wake rake pressures were acquired using PSI (Pressure Systems

Inc.) pressure transducers and the appropriate reference p_tssures and temperatures were recorded. The model pressures were
reduced to coefficient form while the wake distribution of

momentum deficit was integrated to drag coefficient. Reference

4 gives a full description of the wind tunnel and the instrumentation used in the program.

Some of the data from the clean model were used to

evaluate the wall interference in the test section. The results

of that study are contained in-the Appendix; the essential

feature is a correction to attack-angle was necessary due to

lift-induced downwash:

*~ - -0.155

TEST PROGRAM

For the selected blade stations, the operating conditions were furnished by Bell Helicopter for the hover and the

"flat-pitch" (no lift) condition, since these corresponded to

the conditions at which the ice was accreted and at which the

engine performance was recorded. Figure 1 shows the corresponding range of conditions for the stations. It should be noted

that a given station is (ideally) operating at a fixed point on

the plotted lines for the aircraft in hover or in flat-pitch;

actually, due to the finite though low wind speed (9 knots for

flight E) there was a small amount of blade oscillation for this

"near-hover" case. Wind tunnel test points for each model were

selected to be on the indicated operating line for that station

and to bracket the flight test conditions.

Table II contains the list of the conditions for the

tests on the different models along with the measured drag

coefficients; the detailed distributions of the surface pressures and of the wake profiles are contained in reference 4.
RESULTS AND DISCUSSION

In Figures 8 A-D, the drag coefficients from the wind

tunnel program have been plotted for the 4 stations respectively

(Note that C is on a log scale) . The attack angle has been

corrected for the tunnel downwash. Although the data are

sparse, they are sufficiently regular to allow a reasonable

evaluation of the drag change caused by the ice accretions over

the range of the testing. However, additional points would be

useful in forming a better data base. Some typical sets of

pressure distributions are presented in Figures 9A and 9B.

The fact that the very simple simulation of the ice

formations by wooden strips gave drag levels comparable with the

cast replicas is encouraging from the longer-range viewpoint.

The use of simple substitutes of two-dimensional shapes plus

distributed grain roughness is not new (reference 5) and the

results obtained in this test series indicates that, with further research in this area, reliable information may be obtained

by this approach using scale models.

Examination of the pressure distributions (i.e.

Figures 9) shows that the disturbances caused by the ice shapes

were somewhat erratic, but the basic profile still dominated the

field. Differences between the data from the castings and the

strip simulations might be due to resolution (insufficient

number of taps), placement errors on the model, attack angle

differences and deviations between models as well as inexact

modelling. The similarities, however, are again encouraging for

the application, of a modelling technique. Inspection of the

data lead to a conclusion that the changes in lift and pitching

moment would, be small. This was verified by integration of the

pressure coefficient distributions which showed maximum changes
to be within the scatter of the measurements; these were 0.01

for CT and 0.005 for C. over the range of the acquired data.

The, above observations apply to the range of conditions for the test program, which was representative of the

flight range. For example, it would be expected that the ice

formations would produce larger effects outside of the indicated

range of attack angle and Mach number, that the stall angle

would be reduced and that Cr-Max would be less; however the

LI

rotor did .not operate under those conditions.

Overall, the techniques which were applied for the

hover case- have been successful and further work of this type

will provide not only data useful for evaluating helicopter

performance degradations, but a better understanding of the ice

accretion phenomena and of the detailed flow field associated

-with its formations. It is quite possible that some part of the

ice formations may have been lost or eroded during the t.ime

taken to land and document the shapes in many cases; the extreme

detail observed in the molds of flight E in particular would

suggest that that was not the case for that flight. On the

other hand, the relatively close agreement between the data from

the ice replicas and the simple two-dimensional strip simulations would suggest that the fine detail is not important for

the determination of the aerodynamic characteristics of the iced

sections.
APPENDIX: EVALUATION OF TUNNEL INTERFERENCE

Since the rotor section was basically a NACA-001 2

(wit h a chord of 21 inches), there were considerable reference

data available, both from airfoil analysis codes and fro m other

win d tunne l tests; it was one of the primary airfoils used to

determine the magnitud e of the wall interferenc e in both the

OSU/AARL airfoil tunnels [6] and the same methodology was used

in this case.

FEC-channe l 10 is a slotted wall tunnel with a common

plenum , an d henc e downwas h wa s expecte d t o dominat e th e

interference . The procedure was to determine the interference in

subcritical flow and then to separately evaluat e the interactions in supercritical flow.

\

From low speed tests (H =0.3), the local increment (of

pressure coefficient) was determined for specific attack-angles,

i.e. the difference between the pressure coefficients on the

upper and lov/er surfaces adjusted by subtracting the same difference at zero attack angle. This procedure was necessary to

remove the effects of any local surface irregularities in an

otherwise symmetrical model.

These data were then compared with the correct distribution to determine the apparent or free-air attack-angle.

In practice , the "correct" data is obtained from other

"corrected" wind tunnel tests and airfoil analysis codes. The

data may be adjusted to a common Mach number basis, e.g. M=0, by

a Prandtl-Glauert conversion. Once the mean downwash has been

determined, the variations from it may be used to establish the

streamline curvature; in this situation, there were not sufficient data to fully evaluated that part beyond the estimate that

it was negligible compared with downwash.
In the supercritical flov; range, the use of ventilated

walls requires a further analysis since the wall openness determines the location of the transonic terminal shock wave. The

correct shock location (and hence the correc t wall

configuration) can only be determined by testing models of

different chord and/or by referring to "interference-free" data.

In this case, the comparisons showed that the tunnel ventilation

was correct in the range tested; Figures A2 and A3 are typical

of cases where a close comparison could be made, assuming that

the previous subcritical attack-angle adjustment was correct.

Since the downwash was generated by lift, the application of the attack-angle correction is applicable only where the

lift coefficient is linear in alpha. If a fully tapped model

had been available, a proper form for the correction (i.e. in

terms of lift ) would have been possible; however, in the application to this program, the correction is sufficient since

the ranges of application are not large. Also, the specified

correction applies only to a full span airfoil model of chord =

21 inches.

10
REFERENCES

1. Shaw, Robert J. and Richter, G. Paul: The UH-1H Helicopter

Icing Flight Test Program: An Overview. AIAA Paper No. 85-

0338, presented at the AIAA 23rd Aerospace Sciences Meeting,

Reno, Nevada, Jan. 14-17, 1985.

2 . Lee , Joh n D. ; Harding , Rorry ; an d Palko , Richar d L. :

Documentation of Ice Shapes on the Main Rotor of a UH-1H

Helicopter in Hover. NASA CR-168332, Jan. 1984.

3 . Abbott , W. Y . ; Williams , R. A.; Belte , D.; an d

Stella r F. W.: Evaluatio n of UH-1H Hover Performanc e

Degradation Caused by Rotor Icing. USAAEFA Report 82-12,

Aug. 1983. (AD-A141252)

4. Berger, Jack H.; and McDonald , Timothy J.: Win d Tunne l

Tests of Airfoil Shapes Altered by Icing and Airfoil Shapes

with Deicer Boots. FluiDyne Engineering Corporation Report,

Jan., 1984.

5. Bragg , M. B. and Gregorek , G. M . , "Aerodynami c

Characteristics of Airfoils wit h Ic e Accretions, " AIAA

Paper No. 82-0282, presented at the AIAA 20t h Aerospac e

Sciences Meeting, Orlando, Florida,.January 11-14, 1982.

6. Lee, J. D., Gregorek , G. M. and Korkan , K. D. : Testin g

Technique s and Interference Evaluation in the OSU Transonic

Airfoil Facility. AIAA-78-1118, July, 1978.

11
TABLE 1. - SUMMARY OF HOVER ICING FLIGHTS

Flight

Number

Wind

Speed,

A

B

C

D

E

kn

7

10

6

4

Gustiness

Medium

Medium

Medium

Low

9 j Low

LWC

gm/m

0.4

.4

.4

Test

Temp.

C

-12.0

-9.5

-17.5

.4 i -21.5

.7 -19.0

Time In

cloud,

mm.

4.5

4.25

4.0

6.0

3.0

12
JESTS RESULTS

CONFIGURATION RUN # ALPHA SET ALPHA CORR. MACH 1 CD

Clean Wing 1 0.00 0.00 0.313 0.0079

2 0.00 0.00 0.601 0.0079

4 4.00 3.12 0.603 0.0086

5 - 4.00 3.25 0.355 0.0087

6 8.00 6.50 0.355 0.0103

7 8.00 6.25 0.599 0.0210

8 16.00 14.00 0.353 0.0879

9 11.50 8.98 0.603 0.0524

10 11.50 9.34 0.357 0.0145

R-H 11 0.00 0.00 0.358 0.0176

12 5.50 .4.47 0.355 0.0256

13 7.25 5.90 0.355 0.0340

14 10.00 7.80 0.356 0.0802

15 - 2.50 2.00 0.355 0.0195

R-15 16 0.00 0.00 0.423 0.0120

17 2.50 2.00 0.463 0.0120

18 5.50 4.40 0.515 0.0152

19 7.25 5.70 0.543 0.0165

R-19 20 0.00 0.00 0.535 0.0098

21 2.50 1.95 0.593 0.0094

22 5.50 4.23 0.648 0.0121

23 7.30 5.54 0.688 0.0274

Clean Wing 95 0.00 0.00 0.751 0.0080

96 0.00 0.00 0.795 0.0105

97 0.00 0.00 0.894 0.0410

98 1.00 0.75 0.701 0.0080

99 2.00 1.57 0.573 0.0077

100 2.00 1.47 0.746 0.0098

101 2.50 1.95 0.603 0.0078

102 5.00 4.00 0.507 0.0085

103 5.00 3.90 0.600 0.0091

104 5.00 3.77 0.700 0.0192

105 3.50 2.48 0.799 0.0323

107 6.50 5.17 0.519 0.0088

108 6.50 5.08 0.600 0.0104

109 16.00 13.00 0.600 0.1859

F-ii 146 0.00 0.00 0.362 0.0157

147 2.50 2.00 0.351 0.0168

150 10.00 7.87 0.356 0.1177

151 7.50 5.90 0.357 0.0561

152 5.50 4.47 0.355 0.0300

F-15 153 0.00 0.00 0.417 0.0129

154 2.50 2.00 0.465 0.0116

155 5 50 4.38 - 0.512 0.0150

156 'I? 5.74 0.539 0.0205

13
ORIGINAL PASS S3

OF POOR QUALITY

r/R = 0,85

FIGURE 1, PROJECTIONS OF MOLDS OF ICE ACCRETIONS

IN HOVER, FLIGHT E,

14
U.D

C/D

5 0,4

Q_

UJ

& 0,2

X

0

i i i i i i i

/ \

/ \

X

\

\

\

\

i i i i i i j

0,2 0,4 0,6 0.8 1

SPANWISE STATION r/R

FIG, 2 MAXIMUM DEPTH OF ICE ACCRETED ON THE

ROTOR FOR FLIGHT E (HOVER),

15
10

•o

O

•2

O If)

•*>*«—'

x a:

IT)

O «-H

O OO O

• • • *

O 00 O

- •

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• • • •

o o o o

I I

to

a

i

o to ir>

•r- ^ O «—' t—I t-H

**••+ °. °. °.°.

O

 ce

- CD O O O

O III

"Z. O* O «"• <M

O CM CM CM

«*> t—I i—« r-1 t—I

m

OJ UJ

JC •

O —I

•^ E

« e x«*-

ooooooooo

tn

CSILOOunOLOOLO

fn

<

fn

w

a:

E-i

O

CO

o

HI

E-t

u

O

•

ro

w

i-r-f-lt-Hi-i-i—rH

J5 r— •—1«—<t—i r—I f—t r-1»—1»—»

Z <_>««-•••"• •

_ lOO^^^Olf>«»•^Ofvl^-l

B- ooooooooo

. " O"—''—li—It—Ir-It—It—1«—I

16
ORIGINAL PAGE S3

OF POOR OUAUTY

w

D

EH

O

CO

I

M,

vb

so

X

DC

O

3

vo

vo

Cd

QQ

U

Q

M

D

. J

b

•

^

W

17
ORIGINAL PAGE S3

OE POOR QUALITY

w

z

D3

O

u

w

fc,

2

EH

a

EH

(O

CO

w

EH

•J

M

O

CM

«

M

<

•

in

«

a

D

o

M

EM

18
POOR

clean wing

installation

wing with

simulated ice

(RID

FIGURE 6. PHOTOGRAPHS OP MODELS INSTALLED IN FEC CHANNEL 10,

19
CRITICAL

BOUNDARY

-2

0.1 0,5 0.6

MACH NUMBER

FIGURE 7. RANGE OF TEST CONDITIONS FOR THE FOUR SELECTED STATIONS

COVERING HOVER AND FLAT PITCH

20
0,2

0,1

0,02

0,01

0,00^

STN, r/R = O.M

—0-~ CLEAN MODEL

—A—CAST ICE REPLICA

....O«» 2-D WOOD STRIP

•CT

.O0

ATTACK ANGLE (DEG.)

8

.O

10

FIGURE 8. DRAG COEFFICIENTS FOR THE FOUR STATIONS THROUGH THE

TEST CONDITIONS CORRESPONDING TO FIGURE 7 , CLEAN

MODEL , CASTINGS FROM MOLDS AND 2-D STRIP SIMULATIONS

8A RADIAL STATION r/R = 0.44

21
0.10

CD

0,02

0,01

0,004

STN, r/R = 0.60

O CLEAN MODEL

CAST ICE REPLICA

D 2-D WOOD STRIP

0

O

I I

2 4 6

ATTACK ANGLE (DEG.)

8

FIGURE 8. DRAG COEFFICIENTS FOR THE FOUR STATIONS THROUGH THE

TEST CONDITIONS CORRESPONDING TO FIGURE 7 , CLEAN

MODEL , CASTINGS FROM MOLDS AND 2-D STRIP SIMULATIONS

8B RADIAL STATION r/R = 0.60

22
0,10

CD

0,02

0,01

0,004

STN,r/R = 0,77

—O CLEAN MODEL

CAST ICE REPLICA

O

0 2 4

ATTACK ANGLE (DEG.)

STN, 0,93

P-o

0 1

O

i /

FIGURE 8. DRAG COEFFICIENTS FOR THE FOUR STATIONS THROUGH THE

TEST CONDITIONS CORRESPONDING TO FIGURE 7 , CLEAN

MODEL , CASTINGS FROM MOLDS AND 2-D STRIP SIMULATIONS

8C RADIAL STATION r/R =0.77 AND 0.93

23
-1,5

-1,0

-0,5

cr>

00

GO n

UJ U

DC:

o_

0,5

1.0

STATION x/R

M

a =

0,60

0,464

2,0°

— •—CLEAN BLADE

—O—CAST FROM MOLD

--D--2-D SIMULATION

0,1 0,2 0,3

CHORDWISE DISTANCE x/c

0,4

FIGURE 9. PRESSURE DISTRIBUTIONS OVER THE LEADING 0.35 OF THE

BLADE SECTION WITH AND WITHOUT ICE SIMULATIONS

9A RADIAL STATION r/R = 0.60

24
-2.0

-1,5

-1.0

UJ

o

CO

CO

Ijj

-0,5

0

0.5

1,0 -

STATION r/R =0,11

M = 0,355

a = U.U7

•-CLEAN BLADE

CAST FROM MOLD

D--2-D SIMULATION

0 0,1 0,2 0,3

CHORDWISE STATION x/c

FIGURE 9. PRESSURE DISTRIBUTIONS OVER THE LEADING 0.35 OF THE

BLADE SECTION WITH AND WITHOUT ICE SIMULATIONS

9B RADIAL STATION r/R = 0.44

25
3.5

3.0

2.5

I 2.0

Q.

3

CL

I 1-5

CL

O

1.0

0.5

D SET a = 8°, M

CALCULATED, a

CALCULATED, a

0.30

8°

6.5°

Aa =

SET

0.1 0.2

X/C

0.3 0.4

FIGURE Al. DETERMINATION OF THE CORRECTION TO ATTACK ANGLE BY

COMPARISONS OF LIFT-INCREMENT DISTRIBUTIONS

SUBCRITICAL FLOW

26
-3.0

-2.5

-2.0

-1.5

-1.0

-0.5

CP

0

0.5

1.0

Q

O

-X- CRIT

0

O

6

O

LJ

D

O

O FEC, CORR a = 9.0°

M = 0.60, Re = 6.Ox 106

CHORD = 21 in.

OOSU, a = 9.0°

M = 0.60,Re=4.4xl06

CHORD = 6 in.

a

o

8

a

o

a

0.1 0.2

x/c

0.3

D

O

o

0.4

FIGURE A2. COMPARISON OF CP DISTRIBUTIONS IN SUPERCRITICAL FLOW

FEC ATTACK ANGLE CORRECTED FROM FIGURE Al.

( SET ANGLE = 11.5 DEGREES )

27
-3.0

-2.5

-2.0

-1.5

-1.0

-0.5

CP

0

0.5

1.0

O

DD

0

Q

O

CRIT

O

D

O

9

O

D

°o

OFEC, CORR a = 6.24

M = 0.599, Re= 5.0 xlO6

CHORD = 21 in.

DOSU, a = 6.0°

M = 0.598, Re=4.3x10 6

CHORD = 6 in.

a

o a

o o

e

0.1 0.2

X/C

0.3

a

o

0.4

FIGURE A3. COMPARISON OF CP DISTRIBUTIONS IN SUPERCRITICAL FLOW

FEC ATTACK ANGLE CORRECTED FROM FIGURE Al.

( SET ANGLE = 8.0 DEGREES )

28
1 Report No

NASA CR-175089

2 Government Accession No 3. Recipient's Catalog No

4 Title and Subtitle

Wind Tunnel Tests of Rotor Blade Sections With

Replications of Ice Formations Accreted 1n Hover

5 Report Date

March 1986

6 Performing Organization Code

7 Author(s)

John D. Lee, Jack H. Berger, and Timothy J. McDonald

8. Performing Organization Report No

None

10 Work Unit No

9 Performing Organization Name and Address

The Ohio State University

Aeronautical and Astronautlcal Research Laboratory

Columbus, Ohio

11 Contract or Grant No.

NAG 3-374

12 Sponsoring Agency Name and Address

National Aeronautics and Space Administration

Washington, O.C. 20546

13. Type of Report and Period Covered

Contractor Report

14 Sponsoring Agency Code

505-68-11

15 Supplementary Notes

Final report. Project Manager, Robert J. Shaw, Propulsion Systems Division,

NASA Lewis Research Center, Cleveland, Ohio 44135. John D. Lee, The Ohio State

University; Jack H. Berger and Timothy J. McDonald, FlulDyne Engineering

Corporation, Columbus, Ohio.

16 Abstract

Full scale reproductions of 1ce accretions molded during the documentation of a

hover test program were fabricated by means of epoxy castings and used for a

wind-tunnel test program. Surface static pressure distributions were recorded

and used to evaluate 11ft and pitching moment Increments while drag was determined by wake surveys. Through the range of the tests, corresponding to those

conditions encountered 1n hover and 1n flat pitch, Integration of the pressure

distributions showed negligible changes 1n 11ft and 1n pitching moment, but the

drag was significantly Increased.

17 Key Words (Suggested by Author(s))

Helicopter 1c1ng; Airfoil 1c1ng;

Aero-performance degradation;

Wind tunnel testing;

Simulated 1ce shapes

19. Security Classif (of this report)

Unclassified

18. Distribution Statement

Unclassified

STAR Category

20. Security Classif (of this page)

Unclassified

- unlimited

03

21. No of pages

30

22 Price'

A03

*For sale by the National Technical Information Service, Springfield, Virginia 22161
National Aeronautics and

Space Administration

Lewis Research Center

Cleveland. Ohio 44 135

OflktW

Penally tor PiMtt UM <300

SECOND CLASS MAIL

ADDRESS CORRECTION REQUESTED

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National Aeronautics and

Space Administration

NASA-451

NASA