NASA/CRm2000-209921/VOL1

DHC-6 Twin Otter Tailplane Airfoil Section

Testing in the Ohio State University7x10 Wind Tunnel

Dale Hiltner, Michael McKee, Karine La No4, and Gerald Gregorek

Ohio State University, Columbus, Ohio

September 2000

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NASA/CR--2000-209921/VOL1

DHC-6 Twin Otter Tailplane Airfoil Section

Testing in the Ohio State University7x10 Wind Tunnel

Dale Hiltner, Michael McKee, Karine La NoG and Gerald Gregorek

Ohio State University, Columbus, Ohio

Prepared under Grant NAG3--1574

National Aeronautics and

Space Administration

Glenn Research Center

September 2000

Acknowledgments

The authors would like to thank the NASA Glenn Research Center Icing Branch for sponsoring this

wind tunnel test through the grant NAG3-1574 to the Ohio State University. We also would like to

acknowledge the project advocacy and support that the FAA William J. Hughes Technical Center

provided throughout this research effort.

Trade names or manufacturers' names are used in this report for

identification only. This usage does not constitute an official

endorsement, either expressed or implied, by the National

Aeronautics and Space Administration.

NASA Center for Aerospace Information7121 Standard Drive

Hanover, MD 21076Price Code: A06

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Table of Contents

List of Figures ............................................................................................................................... iv

List of Symbols ............................................................................................................................... v

SUMMARY ................................................................................................................................ vi

1. INTRODUCTION ................................................................................................................... 1

2. EXPERIMENTAL FACILITY ............................................................................................. 3

3. MODEL DESCRIPTION ....................................................................................................... 4

4. DATA ACQUISITION .......................................................................................................... 13

5. TEST PROCESS .................................................................................................................... 14

6. DATA REDUCTION ............................................................................................................. 15

7. RESULTS AND DISCUSSION ............................................................................................ 18

7.1 Ice Shape Effects ....................................................................................................... 18

7.2 Velocity Effects .......................................................................................................... 19

7.3 Comparison of Surface Taps and Belt Taps ........................................................... 19

7.4 Repeatability .............................................................................................................. 20

7.5 Section 2 Comparison ............................................................................................... 21

7.6 Flow Visualization ..................................................................................................... 21

7.7 5-Hole Probe Results ................................................................................................. 22

8. CONCLUSIONS .................................................................................................................... 40

Appendix A. Tap Locations ...................................................................................................... 41

Appendix B. Run Log ................................................................................................................ 47

Appendix C. Video Log ............................................................................................................. 49

Appendix D. Solid Wall Correction Calculations ................................................................... 55

Appendix E. Data Reduction Output File Formats ................................................................ 59

Appendix F. DHC-6 Twin Otter Tailplane Coefficient Data ................................................. 63

-) -) ,) ...NASAJCR--,.000-.099.1/VOL 1 m

List of Figures

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

Figure

1 NASA Glenn Research Center Icing Research Aircraft .......................................... 2

20SU Low-Speed Wind Tunnel Facility ..................................................................... 3

3 Section 1 (Baseline) and Section 2 Airfoil Sections ................................................... 5

4 Ice Shape Sections ......................................................................................................... 6

5 DHC-6 Tailplane Wind Tunnel Model Schematic .................................................... 7

6 Model Installation, View from Upstream .................................................................. 8

7 Baseline Installation, _=-26.6 ° (Wake Probe in Foreground) .................................... 9

8 LEWICE Ice Shape Installation (View from Upstream) ......................................... 109 Section 2 Installation .................................................................................................. 11

10 5-Hole Probe ....................................................................................................... 12

11 Pressure Distribution Example .............................................................................. 16

12 Data Reduction Correction Example ..................................................................... 17

13 Axis Systems and Sign Conventions. ................................................................... 22

14 Lift Characteristics, _e=0.0 ...................................................................................... 23

15 Hinge Moment Characteristics, _=0.0 ................................................................... 23

16 Lift Characteristics, Minimum Deflections .......................................................... 24

17 Lift Characteristics, Large Deflections .................................................................. 24

18 Hinge Moment Characteristics, Minimum Deflections ........................................ 25

19 Hinge Moment Characteristics, Large Deflections. ............................................. 25

20 Lift Curve and Hinge Moment Slopes .................................................................... 26

21 Elevator Lift Effectiveness (V=100 kts, Surface Taps). ........................................... 27

22 Elevator Hinge Moment Effectiveness (V=100 kts, Surface Taps) ........................ 27

23 Velocity Effects, Baseline, Lift Characteristics ..................................................... 28

24 Velocity Effects, LEWICE Ice Shape, Lift Characteristics .................................. 28

25 Velocity Effects, Baseline, Hinge Moment Characteristics ................................... 29

26 Velocity Effects, LEWICE Ice Shape, Hinge Moment Characteristics ............... 29

27 Tap Line Comparison, Baseline, Lift Characteristics ........................................... 30

28 Tap Line Comparison, LEWICE Ice Shape, Lift Characteristics ....................... 30

29 Tap Line Comparison, Baseline, Hinge Moment Characteristics ........................ 31

30 Tap Line Comparison, LEWICE Ice Shape, Hinge Moment

Characteristics. ............................................................................................................. 31

31 Baseline Repeatability, V=100 kts, Surface Taps .................................................. 32

32 LEWICE Ice Shape Repeatability, V=100 kts, Surface Taps .............................. 33

33 Section 2 Lift Characteristics Comparison ............................................................ 34

34 Section 2 Hinge Moment Comparison .................................................................... 34

35 Baseline Tuft Flow Visualization, AOA=-8.0 °, _e=0.0 ° .......................................... 35

36 S&C Ice Shape Tuft Flow Visualization, AOA=-8.0°,_=0.0 ° ............................... 36

37 Baseline Cp Distribution, AOA=-8.0 °, 5e=0.0 ° ........................................................ 37

38 S&C Ice Shape Cp Distribution, AOA=-8.0 °, 5e=0.0 ° ............................................ 37

39 5-Hole Probe Dynamic Pressure Comparison ........................................................ 38

40 5-Hole Probe Pressure Differential due to Angle of Attack ................................. 38

41 5-Hole Probe Pressure Differential due to Sideslip .............................................. 39

NASA/CR--2000-209921/VOL 1 iv

List of Symbols

AOA,(x

C

CD

CDo

CH

CHc_

CH8

CL

CLmax

CLc_

CL_

CM,CMI/4

Cp

Ptotal

Pstatic

q,QRe

V

(x,AOA

Of+stall

+e

angle of attackchord

section drag coefficient

section drag coefficient at a = 0o

section hinge moment coefficient

sectmn hinge moment slope versus c_ at constant 6e

sectmn hinge moment slope with 8e at constant

section lift coefficient

peak section lift coefficient

section lift curve slope at constant 8e

section lift coefficient slope with 5e at constant (_

section moment about % chord

pressure coefficient

total pressure

static pressure

dynamic pressure

Reynolds Number

velocity

angle of attack

angle of attack at stall

elevator deflection

NASA/CR--2000-209921/VOL 1 v

SUMMARY

Ice contaminated tailplane stall (ICTS) has been found to be responsible for 16 accidents

with 139 fatalities over the last three decades, and is suspected to have played a role in other

accidents and incidents. The need for fundamental research in this area has been recognized at three

international conferences sponsored by the FAA since 1991. In order to conduct such research, a

joining NASA/FAA Tailplane Icing Program was formed in 1994; the Ohio State University has

played an important role in this effort. The program employs icing tunnel testing, dry wind tunnel

testing, flight testing, and analysis using a six-degrees-of-freedom computer code tailored to this

problem. A central goal is to quantify the effect of tailplane icing on aircraft stability and control to

aid in the analysis of flight test procedures to identify aircraft susceptibility to ICTS. This report

contains the results of testing of a full scale 2D model of a tailplane section of NASA's Icing

Research Aircraft, with and without ice shapes, in an Ohio State wind tunnel in 1994. The results

have been integrated into a comprehensive database of aerodynamic coefficients and stability and

control derivatives that will permit detailed analysis of flight test results with the analytical

computer program. The testing encompassed a full range of angles of attack and elevator

deflections, as well as two velocities to evaluate Reynolds number effects. Lift, drag, pitching

moment, and hinge moment coefficients were obtained. In addition, instrumentation for use during

flight testing was verified to be effective, all components showing acceptable fidelity. Comparison

of clean and iced airfoil results show the ice shapes causing a significant decrease in the magnitude

of CLmax (from -1.3 to --0.64) and associated stall angle (from -18.6 ° to -8.2°). Furthermore, the

ice shapes caused an increase in hinge moment coefficient of approximately 0.02,the change being

markedly abrupt for one of the ice shapes. A noticeable effect of elevator deflection is that

magnitude of the stall angle is decreased for negative (upward) elevator deflections. All these result

are consistent with observed tailplane phenomena, and constitute an effective set of data for

comprehensive analysis of ICTS.

NASA/CR--2000-209921/VOL 1 vi

1. INTRODUCTION

This wind tunnel test is part of the NASA, FAA and The Ohio State University, Tailplane

Icing Program. The purpose of the program is to quantify the effect of tailplane icing on aircraft

stability and control to aid in the development of a flight test procedure that will identify aircraft

susceptible to tailplane icing. As a first step in this program, wind tunnel testing of the

2-Dimensional aerodynamics of the tailplane of the NASA Glenn Research Center Icing

Research Aircraft, a DeHavilland DHC-6 Twin Otter (Fig. 1) was undertaken. Pressure data from

surface taps and belt taps were acquired and integrated to obtain lift, drag, pitching moment, and

hinge moment coefficients. The belt taps consisted of precise holes punched into each tube of a

series of strip-a-tube belts wrapped around the airfoil section. Belt taps were included in the

testing as the aircraft will have belt taps installed during flight testing. Tests were conducted on

the clean, or uniced, airfoil and the airfoil with two representative ice shapes at a full range of

angle of attack (AOA), from large negative to beyond stall, and the full range of elevator

deflections. Reynolds number effects were obtained by testing at two velocities. As shown in

Figure 1, the Twin Otter tailplane is composed of two separate airfoil sections. Basic data for the

clean secondary airfoil section were also acquired. In addition, data were taken from a 5-hole

probe mounted on the airfoil leading edge. This data will be used to determine tailplane AOA

from similar probes in future flight testing.

The representative ice shapes were called S&C (for stability and control) and LEWICE

(because it was generated by the code of the same name). The S&C ice shape has been used in

several stability and control programs on the Twin Otter. This ice shape was determined using a

combination of in-flight photographs of icing on the Twin Otter's tailplane, and the FAA ADS-4

report "Engineering Summary of Airframe Icing Technical Data", by Bowden, et al. The

LEWICE ice shape was predicted using the LEWICE 1.3 computer program. (Specific input

conditions were: velocity = 120 kts, liquid water content = 0.5 g/m 3, median volumetric diameter

= 20_tm, icing time = 45 min, AOA = 0 °, total temperature = -4 ° C.) The ice shape was

determined in five equal time steps. Each step consisted of a 9 minute accretion time, a

redefining of the iced airfoil geometry, and a recalculation of the airflow over the new geometry.The 45 minute accretion time was chosen to be conservative while aligning with the FAA

requirement of demonstrating capability with 45 minute accretion time on an unprotected

surface.

To satisfy program objectives, this 2-D airfoil data will be used to generate separate

aircraft and tailplane stability and control derivatives. These data will allow the tailplane effects

and flow field environment during maneuvering flight to be explored by a flight path simulation

program presently under development. Maneuvers that indicate a tendency for tailplane stall will

be examined analytically refined by flight test. Data obtained from the belt taps and the 5-hole

probes will be used to quantify the tailplane aerodynamics during these flight test maneuvers.

NASA/CR--2000-209921 ]VOL 1 1

'_ _ L Section 2

_ Section 1

<I2._ Irii

Figure I NASA Glenn Research Center Icing Research Aircraft

NASA/CR--2000-209921/VOL 1 2

2. EXPERIMENTAL FACILITY

The Aeronautical and Astronautical Research Laboratory Low-Speed Wind Tunnel

Facility (OSU LSWT) is a closed-loop circuit with a 7x10 ft. high-speed test section, and a

14x16 ft. low-speed test section. A 2000 Hp electric motor provides dynamic pressures up to

65 psf in the 7x10 ft. section. (See Fig. 2.) A sting balance mount allows models to be tested at

angle of attacks ranging from -15 to +45 degrees, while the rotating tunnel floor has a range of

-90 to +270 degrees.

-740 FT

2m

6 BLADE FAN - 20 FT.

I" 1I I

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SCREEN TEST SECTIONI

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WALLS

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Figure 20SU Low-Speed Wind Tunnel Facility

Tunnel speed is manually set based on average values obtained from two dedicated tunnel

pitot and static probe sets. A PC-class computer is used to control sting and mounting table

angles, and wake probe sweep, as well as data acquisition tasks. Up to 14 data channels can be

sampled. Signal conditioning and amplification are provided by individual rack mounted units.

Testing of 2-dimensional airfoils is conducted by installing a 7-ft. span model to fill the

height of the low speed test section. The model is attached to the floor mounting table, which is

then rotated to provide specific AOA's. Surface taps on the model are used to acquire pressure

distributions, which are integrated to obtain model forces and coefficients. To determine drag,

wake survey probes are installed down stream of the model. By sweeping these probes across the

test section, the complete model wake can be measured.

NASA/CR--2000-209921/VOL 1 3

3. MODEL DESCRIPTION

The two airfoil sections tested are shown in Figure 3. As is seen from the 2-view aircraft

drawing (Fig. 1), the Section 1 airfoil was a significant portion of the tailplane area. Drawings of

the two ice shapes are shown in Figure 4. Some minor smoothing of the lofted airfoil shape was

required; a 0.020 inch "bump" at the lower leading edge was removed, and a straight line contour

from the elevator main spar to the trailing edge was assumed.

The primary factors in the model design were to provide a smooth aerodynamic surface

while allowing flexibility in model configuration and low fabrication costs. A highly stiff design

was also desired, to maintain model shape under load. The chosen model design consisted of

airfoil sections machined in machineable plastic (also known as REN material), supported by

metal tubing spars and aluminum fibs. This allowed CNC machining to be used to fabricate the

airfoil shape with high accuracy.

A schematic of the model is shown in Figure 5. The 20 inch span Machineable Plastic

sections were supported by two internal metal tubing spars on each of the stabilizer and elevator

sections. Three fibs and two endplates structurally connected the spar tubes of the stabilizer to

the forward spar tube of the elevator. The ribs contained keyways for keys mounted on the

stabilizer spar tubes to provide stiffness under model moment loading. The ribs of the elevator

overlapped those of the stabilizer, to form a hinge around the forward elevator spar tube. End ribs

on the elevator provided attach points to the endplates, which were used for changing the

elevator deflection angle. The machineable plastic sections transmitted their loads to the ribs

through pins. The endplates contained rods to attach the model to the tunnel yaw table floor

supports, and to a ceiling pivot. The model was mounted to the yaw table such that the tablecenter of rotation was at 28% model chord.

Removable plug sections were used to create the two configurations tested. The stabilizer

section consisted of the baseline (or Section 1) section. An aft extension plug was then used to

create the Section 2 stabilizer section. The elevator section consisted of the Section 2 section

with plugs to create the baseline section. Figures 6 through 9 show the model as installed in thetunnel.

Tap locations were chosen to provide high resolution where pressure gradients were

large, and low resolution elsewhere to minimize the total number of taps. Taps were labeled

starting with #1 on the leading edge of the stabilizer and ending with #86 on the elevator trailing

edge. (This included taps for the baseline and Section 2 airfoils.) Odd numbered taps were

located on the suction side of the airfoil, with even numbers on the pressure side, except at the

boat-tail ends. The tap locations of the ice shapes replaced those on the nominal airfoil, which

was reflected in their numbering. The final tap location configurations are listed in Appendix A,

which also shows a graphic of their positions on the airfoil sections.

Surface taps were installed in-house prior to the completion of the section assembly. Taps

were located on a 12-inch airfoil section using a Bridgeport milling machine with digital distance

readout. Tap locations were determined relative to a coordinate system readily indicated using

available points on the section. Once the locations were marked along the section moldline,

0.0625-inch holes were drilled normal to the surface 0.5 inches below one edge. Mating holes

were then drilled vertically into the milled out end of the section. Stainless steel tubing was then

fitted to these mating holes and routed along the section end and through tubing passageway

holes. The tubing length was sized so at least 2 inches extended beyond the section. The tubing

was glued and sealed to the mating hole using plumbers paste. With the tubing completed, the

NASA/CR--2000-209921/VOL l 4

milled out end was filled and glued to a mating 8 inch section to complete the 20-inch section.

Strip-a-tube tubing was then attached to the stainless steel tubing and routed through the

passageways in each section to connect the surface taps to the scanivalve ports. The tubing length

was approximately 15 ft.

The belt was installed by first milling a slot into the pressure side of the stabilizer and

elevator. Double sided tape and epoxy were then used to secure the belts to the section. The free

ends of the belts were then routed through the slot and into the tubing passageways. Tap

locations were determined by measuring a surface distance along each side of the belt. A special

tap tool was then used to punch a hole in each tube and remove the excess material. Note that

each belt was continuous, connecting directly to the scanivalve, and only one tap per belt tube

was used. The two outer tubes of the full belt did not contain any taps, since experience has

shown that local flow accelerations over these outer tubes cause erroneous pressure

measurements.

The 5-Hole Probe is shown in Figure 10. As indicated in the drawing, the probe provides

six pressures. Ptotal and P_ta,ic are used to determine velocity. The difference between the upper

and lower static port pressures in the probe head provides an indication of AOA, and the

difference between left and right static port pressures provides an indication of sideslip angle.

The probe was attached to the leading edge of the model with a "glove" fixture that

wrapped around the leading edge. The alignment tabs of the probe fit into slots in the probe

mount on the glove. A set screw in this mount locked the probe in place. The glove was secured

to the model by a screw in each comer. The probe pressure tubing was routed through a hole

drilled through the model into the front pressure tubing passageway. The probe was mounted at

approximately 16 inches above the tunnel floor, at 75% of the lowest model span section. Note

that this probe installation duplicated the installation that will be used on the Twin Otter for

flight testing.

Section 1 (Baseline)

Section 2

Figure 3 Section 1 (Baseline) and Section 2 Airfoil Sections

NASA/CR--2000-209921/VOL 1 5

LEWICE]ce$hope

1"4

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-2

JJ

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Figure 4 Ice Shape Sections

iiii

3

NASA/CR--2000-209921/VOL 1 6

End Prate

(1 of 2)

Main Spar

_ Tube

I

[ Rib (1 of 3)

Elevator

Plug Section

(1 of 4)

Spar TubeI Installation

I (4/Rib)

II

Pin Hole

Elevator

Section

(1 of 4)

t II II II II

Pressure

Tubing

Passage

(3/Rib)

Elevator

HingeLocation

Stabilizor

Section

(1 of 4)

Elevator

Deflection

PositioningRib (1 of 2)

MountingTube

Figure 5 DHC-6 Tailplane Wind Tunnel Model Schematic

NASA/CR--2000-209921/VOL 1 7

Figure 6 Model Installation, View from Upstream

NASA/CR--2000-209921/VOL 1 8

Figure 7 Baseline Installation, 5e=-26.6 ° (Wake Probe in Foreground)

NASA/CR--2000-209921/VOL 1 9

Figure 8 LEWICE Ice Shape Installation (View from Upstream)

NASA]CR--2000-209921/VOL 1 10

Figure 9 Section 2 Installation

NASA/CR--2000-209921/VOL 1 11

NOTES

t.tV, Z$(NIL C0

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_,4n Pipa_ "ll.4eF _ :Sdi_r/J_o_ _'popAll i[_<X,il=._ TT'.?i=, )_&ilr't pl,_,_ ]l'_l,a frO_,l.

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NOTESl. PPcSarl_ L_4_, "1;_._ (_ _il f_01tlpl'p$ P_'Pk,

a.-b-. _,',,, lo IIr."l,'_ _ TUB[ _1 I_T_

3. g ll_airl i Proof-... l_",a,_¢r$_oc-I J. V_. JL'3'3.-EI4_' g k,_25 L_.

far f,r,r .P,_a_,'fm I]6,e_lia_

Figure lO 5-Hole Probe

NASA/CR--2000-209921/VOL 1 12

4. DATA ACQUISITION

The scanivalves contained 48 ports and a single Statham 2.5 psi differential transducer,

vented to atmospheric pressure. The individual ports were accessed by mechanically rotating a

perforated ring to individually pass each port pressure to the transducer. This action was

controlled by computer, allowing up to 3 seconds between samples of each port. Four units were

used providing up to 192 pressure ports. To provide higher accuracy, tunnel total and static

pressures were applied to the first four ports of each unit. During data reduction, these pressures

were used to provide tunnel conditions for the pressure coefficients obtained from each

respective unit.The wake measurement system consisted of pressure probes mounted to a 16 ft. length of

aerodynamic tubing which spanned the test section. The tubing was supported by roller guides

located at each wall exterior, and one roller guide mounted on a vertical section of aerodynamic

tubing located at approximately 1/3 of the tunnel width. Three sets of one total and one static

pressure probes were attached to the aerodynamic tubing. These probes were 4 inches apart with32 inches between sets to allow some overlap during the 36-inch sweep across the test section.

An individual probe was used for each total and static pressure, allowing measurements to be

made at a constant tunnel cross-section location. Both total and static wake probe pressures were

measured by 2.5 psi differential transducers.Tunnel conditions were measured using dedicated 2.0 psi differential transducers for pitot

pressures, while 0.5 psi differential transducers were used for the static pressures. All differential

transducers were vented to the atmosphere, with atmospheric pressure determined from a

sensitive altimeter.

All transducers were calibrated using a manually operated water manometer. Calibrations

of the scanivalve transducers over both positive and negative pressure ranges were linear within

0.32% full scale, with a standard deviation within 0.15%. Repeatability of the calibration slopes

was within 0.5% full scale. The wake probe transducers were linear within 0.1% with standard

deviation less than 0.03%. For the tunnel transducers, linearity was within 0.1% and standard

deviation within 0.03%. Similarly, tunnel static transducers were calibrated to 0.2% and 0.12%,

respectively.Power for all transducers was supplied by signal conditioners mounted in the control

room. Amplifier banks mounted in the control room amplified the transducer outputs to the

proper voltage for the 14 channel A/D converter. The digital outputs were then fed to a 386 classPC. This PC contained all data gathering and data reduction programming, with only tunnel

airspeed being manually controlled. Angle of attack positioning, operation of the scanivalves,

operation of the wake probe sweep, and recording of all raw data was performed by the run

program. During an AOA sweep, the run program displayed current tunnel conditions and model

position, as well as the value of each active A/D channel.

NASA/CR--2000-209921/VOL 1 13

5. TEST PROCESS

Appendix B contains a Run Log for the test. (Note that runs prior to run 10 did not

generate valid pressure data due to a malfunctioning scanivalve.) Note that repeatability runs of

the baseline configuration at 100 kts were taken at Be=0.0 °, 14.2 ° and -26.6 °. Also, the last data

point of each run was taken at AOA=0.0, as a run repeatability check. Only a few runs weretaken with the Section 2 configuration, due to test time limitations.

The daily test procedure consisted of first adjusting all signal conditioners and amplifiers

to their proper values. This compensated for the small drift associated with the particular

electronics used. The barometric pressure in the "balance room" was noted each morning. (All

differential transducers were vented to the "balance room" atmosphere.) After the tunnel

inspection was completed, the run software was started.

The run software was initially set up for the model characteristics prior to any testing.

This included model reference dimensions, as well as any transducer calibrations. The "balance

room" pressure and the tunnel temperature were entered each morning. Other specific items

entered prior to each run included; run number, run description, the specific set of pressure taps

in use, elevator deflection angle, the file name into which the raw data would be saved, status of

wake probe usage, and specific AOAs desired during the run. Once these were set, a wind off

tare was taken. This numerically zeroed all the pressure transducers to ambient conditions, within10 -3 psid.

With the tares completed, the run program was started which positioned the model to the

first AOA of the sweep. The tunnel fan was then started and manually set to a specific RPM,

which resulted in the desired dynamic pressure, as displayed by the run program. When this

dynamic pressure was stabilized, the AOA sweep mode was engaged. The run program then

positioned the model and took data for all desired AOA points. A typical AOA sweep of

10 points took approximately 30 minutes to complete. A wake probe sweep, if run, occurred after

pressure tap data were taken at each AOA. This sweep added an extra 2 minutes to the time to

take data at each AOA. Because of this, only a limited number of wake probe sweeps wereaccomplished.

Note that once the RPM was set, it was not changed during a run. Large blockage at the

higher AOAs caused a reduction in the tunnel velocity and Reynolds Number. Nominally this

reduction was on the order of 10 kts (Aq=6 psf and ARe=0.5xl06), during the 100 kts runs. For

the runs into the AOA=20 ° range, the reduction was on the order of 20 kts (Aq=l 1 psf and

ARe=0.8xl06). For the 60 kts runs, the reduction was typically less than 10 kts (Aq=3.5 psf and

2uRe=0.45x 106), with 15 kts (Aq=4.5 psf and _uRe---0.55x 10 6) being an extreme. This reduction in

velocity was allowed because the time required to manually set the velocity at each AOA point

would have been excessive. Also, the incremental velocity between each 60 and 100 kts run was

kept within a reasonable range, so any Reynolds number effects would be apparent.

With the AOA sweep completed, any desired post run note was entered into the run

program, and the raw data were saved to the hard disc. The raw data were the final result

generated by the run program. The raw data file included the output voltages of all transducers,

and the tap location and calibration data. This method of data storage allowed a run to be

modified post-test if errors in the setup were discovered after testing was completed.

NASA/CR---2000-209921/VOL 1 14

Limited tuft flow visualization was also performed. A single row of tufts was installed in

the center of the upper and lower model sections. These tufts were briefly video taped at each

AOA during most of the AOA sweeps. Correlation to the specific test condition was

accomplished by logging the time of each recording. Appendix C contains the timing log for the

video. (Note that runs 01 through 07 were applicable for the tuft video.) Additionally, some runs

were made with more complete tufting of the model to check for 3-Dimensional effects. Pictures

and video were taken of most of these runs. (These runs are indicated as Flow Viz runs in the

Run Log of Appendix B.)

6. DATA REDUCTION

Coefficient data were obtained by integrating the pressure coefficient distributions of the

stabilizer and elevator separately. A typical pressure distribution is shown in Figure 11. This was

accomplished by applying the averaged pressure of adjacent taps to the area between them. When

summed, these values gave coefficients defined relative to the local element coordinate system.

These local coefficients were then converted to a lift and drag contribution by the appropriate

axis transformation. Moment was obtained similarly, using moment arms of the distance from a

reference center to the center location of each integration area.

The solid wall correction factors discussed below use drag coefficient as a primary

parameter. Due to the limited amount of wake data obtained during the test, all drag coefficient

values were estimated from pressure drag coefficients. A Coo bias coefficient was determined by

comparing wake drag coefficients to pressure drag coefficients at AOA=0 °, and 5e=0 °, for each

Section and ice shape configuration tested. These Coo coefficients were then applied to all

pressure drag coefficients obtained with the same Section and ice shape configuration. The

resulting drag coefficient values were judged to be well within required tolerances for obtaining

accurate solid wall corrected coefficients.

With all coefficients thus obtained, solid wall corrections were then applied. The

formulas used to obtain the corrected data are shown in Appendix D. An example of typical

magnitudes of the corrections is given in Figure 12, which shows corrected and uncorrected data

for a baseline run. Note that a buoyancy term was not included in the corrections, which is used

in Ref. D1 as a CD correction. Ref. D1 implies that a buoyancy correction is valid only for large

elements of an aircraft configuration, and is not significant for airfoils. Sample calculations

verified this implication.

In addition to the corrected coefficient data, the data reduction program also calculated

other important data. Five output files were generated containing uncorrected, corrected, wake

drag, 5-hole probe pressure, and wall static pressure coefficients. Appendix E contains a

description of the file name and data structure conventions of the output files.

NASA/CR--2000-209921/VOL 1 15

DHC-6 Taitplane OSU 7X1 0 Wind Tunnel DataRun 55, AOA=-8.0, Uncorrected Data

c-

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Figure 11 Pressure Distribution Example

NASA/CR--2000-209921/VOL 1 16

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DHC-6 Tailplane OSU 7X1 0 Wind Tunnel DataRun 55, Uncorrected/Corrected Comparison

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Figure 12 Data Reduction Correction Example

NASA/CR--2000-209921/VOL 1 17

7. RESULTS AND DISCUSSION

The purpose of this test was to quantify the effects of ice shapes on the Twin Otter

Tailplane 2-dimensional aerodynamics. This quantification was sought relative to the tailplane

stall phenomenon. The two most important phenomena were therefore lift and hinge moment.

These are discussed in the following section. Appendix F contains a complete data presentation

of all test results of CL, CD, CM, and CH.

All data are presented with the tailplane as installed on the aircraft, using an aircraft

reference system. Therefore, when the airfoil is generating lift relative to the airfoil reference

system, this corresponds to negative lift, or positive tailplane down force, in the aircraft reference

system. Other sign conventions are AOA positive with the stabilizer leading edge up, moment

and hinge moment positive when nose up, and elevator deflection positive with trailing edge

down (causing positive lift). Figure 13 presents a schematic describing the axis system and signconventions.

7.1 Ice Shape Effects

The nominal lift characteristics with zero elevator deflection are shown in Figure 14.

(Note that the CL at AOA=-21 ° of the baseline configuration was judged to be in error, and

indicative of the largely separated flow at this AOA. A more appropriate CL value was estimated

to be CL=-1.22, based on the characteristics of the 60 kts data.) Here the significant reduction in

negative CLmax and decrease in negative AOA at stall due to the ice shapes is clearly seen.

Specifically, negative CLma× is reduced from CLmax=-l.3 for the baseline airfoil, to CLmax=-0.64

for the S&C ice shape, and to CLmax=-0.60 for the LEWICE ice shape. Negative AOA at stall is

decreased from the baseline's AOA=-18.6 ° to approximately AOA=-8.2 ° for either ice shape.

Nominal hinge moment characteristics are shown in Figure 15. These clearly show the ice shapes

causing an approximate ACH=0.02 increase at stall. Note, for the LEWICE ice shape, the

transition is not as abrupt as for the S&C ice shape.

The effects of the ice shapes on elevator effectiveness are shown in Figures 16 through

22. (Only the LEWICE ice shape data is shown, as there is little difference in the results with

either ice shape.) Figure 20 shows that, in the low AOA linear range used to determine the lift

curve slopes, the uniced and LEWICE airfoils have similar trends, with both slopes peaking at

8e=10 °. A noticeable effect of elevator deflection is that negative AOA at stall is decreased for

negative elevator deflections and increased for positive tail deflections. For LEWICE data

(Fig. 16) at _Se=-10°, CLm_x=-0.85 at AOA=-6.25 °, while at 5e=+10 °, CLmax=-0.33 at AOA=-10.2 °.

These trends continue when the negative deflection is increased to 5e=-20 ° resulting in

CLmax=-l.21 at AOA=-4.32 ° (Fig. 17). However, increasing positive deflections result in a

smaller increment for CLmax, of ACL=-0.18 and no change in AOA at stall (Fig. 17). Figure 21

shows that the elevator lift effectiveness is extremely low beyond _=-20 °. This indicates that the

elevator is saturated, that is, in fully separated flow, at the larger negative deflections.

These stall AOA and effects of flow separation are also seen in the hinge moment, shown

in Figures 18 through 20. There is only a small change in CH due to the LEWICE ice shape

between _5_=+10 ° and _5e=+14.2 °. However, for both deflections there is a significant shift in CH

starting at AOA=-8 °. This shift is significantly trailing edge down, of roughly ACH=+0.040, and

is close to the clean 5e=0.0 ° values at AOA=-12 °. The effect of ice on the 8_=-10 ° data shows

only a small increase in CH, at a post stall AOA. However, both the iced and uniced data show a

negative steepening of the CH curve at this deflection. At _ie=-20 ° there is a significant ACH=0.02

NASA/CR--2000-209921/VOL 1 18

increase,startingat anegativeAOA belowstall of AOA=-2.3°. This results in the sharp negative

increase in CH6 for deflections between _Se=-10 ° and 8e=-20 ° shown in Figure 22. For other

deflections, the CH_ for the LEWICE ice shape shows similar trends to that for the uniced airfoil.

Figure 20 shows that, while CH,_ of the baseline airfoil is negative and relatively constant, CH,_ for

the LEWICE ice shape shows a trend of negatively increasing with increasing positive or

negative elevator deflection.

These results match well with observed tailplane stall phenomena. During approach, the

elevator would most likely be at a negative deflection. Tailplane stall could more readily occur,

since the AOA at which stall would occur would be lowered. Once the stall occurred the AOA

would become more negative due to the nose down pitching dynamics. With a negative CHc,, the

trailing edge down (positive) hinge moment would increase, keeping the elevator in a trailing

edge down position. The elevator would continue its trailing edge down tendency, until it reached

a position limit or the AOA stabilized. Also, in general, the trend in CH6 is decreasing with

increasing positive tail deflection. This would support the typically observed "stick force

lightening" phenomenon. However, the trend in CH6 at the maximum positive tail deflections and

largest negative AOA is significantly increasing (Fig. 22). This suggests that any "stick force

lightening" would be dependent on the specific AOA and deflection of the maneuvering aircraft.

7.2 Velocity Effects

Representative velocity effects are shown in Figures 23 through 26, for the baseline and

LEWlCE configurations. Figure 23 indicates that for _5_=0.0 ° there is a AAOA=2 ° decrease in

negative AOA at stall at 60 kts (Re=2.7xl06) compared to the 100 kts (Re=4.8xl06) data. There

is also an earlier stall break at the positive AOAs for the +Se case. In general, there is an increase

in negative lift at the higher speed. At 5_=-20 ° there is a significant decrease in negative CL at

about AOA=-4.0 °, which is most likely caused by the largely separated flow of this deflection.

These results are consistent with typical Reynolds number effects.

Figure 25 shows an indication of the earlier stall break for the Be=0.0 ° case in the hinge

moment data. Some slight increase in hinge moment is seen for the lower speed, but this is

nominally in the repeatability range and so is not considered to be significant.

The LEWICE ice shape data of Figures 24 and 26 show little differences due to the speed

variation. Except for some noticeable larger differences at the most positive AOA, such

differences that are seen are within nominal repeatability.

7.3 Comparison of Surface Taps and Belt Taps

In general the agreement between the surface tap and belt tap results can be qualified as

good. However, significant differences seem to occur at conditions of largely separated flow.

Figures 27 and 28 show lift characteristics for the baseline and LEWICE cases. Agreement is

good for the baseline data, until near stall at 8_=0.0 °, and at the low negative AOAs for _e=-20 °.

In the later case, the shift is roughly ACE=0.2, which is relatively large. (Note, however, from

Appendix F, Figs. F.27 and F.28, that the drag at low AOA seems to be erroneously large.

Applying solid wall corrections using the drag from the surface taps reduces this shift to

ACE=0.1, which is still significantly large.) Also, a reduction in negative AOA at stall of about

AAOA=2 ° is seen at _Se=0° and negative 5_s. The LEWlCE ice shape data shows a lowering of

negative CL of about ACE=0.06 at 5_=-20 °, and a slight increase in negative CL of ACE=-0.05 post

stall for _=0 °. Note that the stall AOAs agree.

NASA/CR--2000-209921/VOL 1 19

The hinge momentdata of Figures29 and 30 show someshifts in CH at 5e=0.0° and_5e=-20° for the baselinecase.The ACH=0.04 shift with _ie=-20 ° near AOA=-4.0 ° is large,

however, agreement is good near the stall at roughly AOA=-14 °. The constant shift of the

baseline data with _=0 ° of about ACH=0.015 is significant, but the trends are more closely

followed. There is no indication of a lowering of the negative AOA at stall for the belt data. The

LEWICE data follows the trends of the baseline data, but the differences are of much smaller

magnitude.

Overall, as tested, the belt is considered to be a good data source for predicting trends.

However, care must be taken in using the belt results. The trend of the better agreement of the ice

shape data is encouraging. However, for the clean airfoil, the lowering of the negative stall AOA

and the shift in CH and negative 8_ must be taken into account. Also, there are noticeable belt

effects starting at approximately AOA=-4 ° with 8_=-20 ° for both the iced and uniced airfoils.

These effects indicate that the belt data is more suspect in areas of large flow separation, possibly

due to span-wise flow over the belt. As such span-wise flows are more likely to occur on the

Twin Otter tailplane in flight, the belt data could be more significantly effected. Therefore, the

flight test data obtained from the belt will need to be carefully evaluated.

7.4 Repeatability

In order to quantify repeatability of the coefficient data, repeat runs were made at 100 kts

and Be=0.0 °, _5e=-26.6 °, and 5_=14.2 °. Except for a few AOAs where there is significant

separation, the repeatability is good. Figures 31 and 32 show repeatability comparisons for the

baseline and LEWlCE ice shape.

From these figures it is seen that nominal differences in CL for the baseline case using

surface taps are within ACL=+/-0.03 and ACH=+/-0.007. At some extremes for the _5_=-26.6 ° data,

the differences are within ACL=+/-0.1 and ACH=+/-0.03. The LEWICE ice shape data agreement

using surface taps was better than that for the baseline case. Nominal differences in CL are within

ACL=+/-0.02 with extremes to ACL=+/-0.04. Nominal differences in CH repeatability are within

ACH=+/-0.007. All are good values.

In order to quantify more precisely these levels of repeatability, standard deviation values

were calculated for all corrected coefficients obtained with surface taps at each target AOA. (No

correction to any coefficient was made for shifts in AOA away from the target values.) Nominal

standard deviation for CL was judged to be within CL=0.015. This nominal value neglected about

10% of the data points, which was typical for the assessment of all nominal values. The largest

standard deviation value was CL=0.18, for the baseline airfoil at target AOA= 12.0 ° and 8_= 14.2 °.

Nominal standard deviation for CD was Co=0.05, with extremes to CD=0.10. Both CM and CH

had nominal standard deviations of 0.003, with extremes to CM=0.03, and CH=0.21 at target

AOA=12.0 ° and _5e=14.2 ° for the baseline airfoil. Applying the 2-sigma level of uncertainty

resulted in repeatability within ACL=+/-0.03, ACD=+/-0.01, ACM=+/-0.006, and ACH=+/-0.006.

These are considered to be good overall repeatability levels, which also agree with the above

assessment obtained by viewing the repeatability plots.

Repeatability for the baseline configuration using belt taps showed similar nominal

values. However, at the extremes, the differences in coefficient values could be up to 3 times

larger than those for the surface data. Repeatability of the LEWICE data using belt taps was the

same as that of the surface tap data.

NASA/CR--2000-209921/VOL 1 20

Also, note that, due to model twist under load, it is judged that AOA is measuredtowithin AAOA=+/-0.5 °. This size of error, along with the size of coefficient errors defined above,

results in the symbol size of the plots of Appendix F representing the approximate nominal error

range of the data.

7.5 Section 2 Comparison

The Section 2 airfoil was tested at 60 kts with _Se=0.0°, 14.2 °, and -20.0 °. The 5-Hole

probe was not installed during this testing, as it had been removed earlier with no change in the

baseline results. (No other deflections or ice shapes were tested due to time limitations.) Lift and

hinge moment comparison plots are shown in Figures 33 and 34.

The most distinctive characteristics of the Section 2 airfoil, compared to the baseline, are

its higher CLmax=-l.36, approximately AAOA=2.0 ° lower negative AOA at stall, and the

sharpness of the stall break. Section 2 has a longer chord for the main element, significantly

smaller gap, and no extension of the elevator leading edge. All of these configuration differences

contribute to this higher CLma×. The hinge moment behaves as expected, with a larger CHs, caused

by the lack of the moment balancing elevator extension of the baseline elevator.

7.6 Flow Visualization

Additional tufts were added to the model at run 60 to provide more complete flow

visualization. Overall, 2-Dimensional flow quality can be characterized as good for all airfoil

configurations. For the baseline with 5e=0.0 ° the tufts showed some spanwise flow towards

model centerline near the tunnel floor and ceiling due to the tunnel boundary layer interference.

Additionally, a slight flow angle away from the ribs towards the centerline of the tapped airfoil

sections was seen. This could be described as a "rib interference" phenomenon. Also, in general,

the flow over the belt section appeared to separate later than the surface tap section. The most

significant flow field phenomenon observed with 8e=-26.6 ° was separated flow on the elevator at

all AOAs tested. The S&C ice shape with 5e=0.0 ° and 8_=-26.6 ° showed little 3-Dimensional

flow effects. The "rib interference" phenomenon was not as apparent, and separation of the

elevator occurred along a constant chordline.

Comparing the observed flowfield to the surface to belt comparisons of c) above, shows

that most of the flowfield phenomena, while visible, did not cause significant differences in

coefficients. Baseline coefficient data showed significant differences only when the flowfield

was completely separated. This was most noticeable at 8e=-26.6 ° and post stall.

Figures 35 and 36 show examples of tuft photographs for the baseline and S&C ice shape

at AOA=-8.0 ° and V--100 kts with _e=0.0 °. The corresponding Cp distributions are shown in

Figures 37 and 38. The figures show clearly the large amount of separation that occurs with the

ice shapes at low AOA. In general, they also show the constant chord line of the flow separation

across the span, except near the tunnel floor and ceiling where boundary layer interference has an

influence on the flow. The separation line of the ice shape is indicated in the Cp distribution plot

to be at approximately 40% chord, which has fully enveloped the tufts at approximately 45%

chord. (Note that the line of tufts near the ceiling is judged to be influenced by the ceiling

boundary layer interference, and so is not indicative of the actual 2-D airfoil stall location.) The

characteristic bluntness of the Cp distribution caused by the ice shape is clearly evident. The

baseline airfoil has attached flow over most of its chord, with only beginnings of separation

indicated near the trailing edge. This is indicated in the Figure 37 by the classic shape of the Cp

distribution, except where the start of separation causes a flattening of the upper surface

NASA/CR--2000-209921/VOL 1 21

pressuresstartingat approximately80%chord. Overall, the tuff flow visualizationshowswelldefined2-D flow andgoodagreementwith theCpdistributions.

7.75-Hole Probe Results

Plots of the 5-Hole probe data from the 100 kts baseline run are shown in Figures 39

through 41. In Figure 39, the dynamic pressure indication of the probe is shown to be lower than

tunnel conditions by approximately 6 to 10 kts. (Note, the indicated runs were within +/-10 kts of

the target of 100 kts.) Also, there is a tendency for lower indications of dynamic pressure at the

larger positive and negative AOAs tested. The AOA ports show good fidelity in indicating AOA

(Fig. 40). Assuming a reasonably linear probe calibration, the reduction in upwash angle that

occurs when the iced airfoil is stalled is clearly seen. The sideslip ports show a small constant

bias in Figure 41, which may be due to a slight tunnel angularity, or the probe itself.

The figures also indicate differences in the probe response with airfoil configuration. This

suggests that each airfoil configuration will need its own probe calibration. However, the

magnitude of the noted differences on the probe output values is currently not known, as a probe

calibration has not been accomplished. An assessment of all aspects of the data fidelity of the

probe will be accomplished when the probe calibration is available.

/1+ Pitching

Moment

J

+ Angle-of-Attack

+ Lift

1'

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( + Downforce)

Figure 13 Axis Systems and Sign Conventions

NASA/CR--2000-209921/VOL 1 22

DHC-6 Tailplane OSU 7X1 0 Wind Tunnel DataBaselTne V=7 O0 kts Surface Taps

d

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Figure 14 Lift Characteristics, _e=0o0

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DHC-6 Tailplane OSU 7X1 0 Wind Tunnel DataBaseline V=100 kts Surface Taps

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Figure 15 Hinge Moment Characteristics, 5e=0.0

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: ---A--- S&:C Ice Shape

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NAS A/CR--2000-209921/VOL 1 23

DHC-6 Tailplane OSU 7X1 0 Wind Tunnel DataV=I O0 kts Surface Taps

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Figure 16 Lift Characteristics, Minimum Deflections

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Figure 17 Lift Characteristics, Large Deflections

NASA/CR--2000-209921/VOL 1 24

DHC-6Tailelane OSU 7XlOWTndTunne_Data

V=lO0 kts Surface Taps

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Figure 18 Hinge Moment Characteristics, Minimum Deflections

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Figure 19 Hinge Moment Characteristics, Large Deflections

NASA/CR--2000-209921/VOL 1 25

V=lOOkts Surface taps

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Figure 20 Lift Curve and Hinge Moment Slopes

NAS,adCR--2000-209921/VOL 1 26

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0.00 : " : " : : ...... : : ......

0.08

0.05

_" 0.04

3...... iiilli__ii' i i i' !.i/li _I(',i_;_' ....

0.02 ....

0.00 : : : ' : : : ....... :':::' ;;';:i-18 -14 -10 -6 -2 2 6 10

AOA (deg)

0.08

0.08

o.o_

0.02

e_mi._-2o-ze J

":::1: ::1: ::1 :: :1: : :1: ::": :i]0.00 ................

-18 -1¢ -10 -6 -2 2 5 10

^o_ ((_)

Figure 21 Elevator Lift Effectiveness (V= 100 kts, Surface Taps)

0000

-0002

-0004-

-0006

-0008

0.0o0

¸i¸.i.14.1¸1-i[-i i i l i.i i.lli I.L._,-I.I :

i i ii i ii_0 _002

i,-_' -i-/i i i"-:, i! i _

::i:i:l;_i:i::l_::i:::l:i__-_;:-i- _ -0004-....

:: ......... : ::_I i:!_'i/!!i_!:_,ii : :ii 'i: !:: -o.oo6:.

_,:,_,_E%}%.... . : i :i i -, ! !i-

-0.008

.: " I " : ] : : ] :..:.; L: .: :.I .: : .:.J :.: .:.

ii ........... iiiiiililtliliill!! iiti!i_ili-: :-;-r-.-;-:--I-;-:-:-i ;-:-; I--: .:-:--I -: -. 1 ;-. -:-

i -i: \ii- i-:- -!-i "%-i

.: : : .... 2.:..I.:.:..:.] :..: : L : 2.2 I : :-.J :-: :

: : : : ' : : ;" ; , ,

O00Q

-0 002

-0004

-0 006

-0008-18

/'.!2 :.!.i ',.JL.!..].!!.!.IL!!.',.!L:.:.:.:.:..

.:..:.:..Li.: ..: :.:.. ;.:..:/.:.: ;..:.: ;.--."::_-.:..

+i_/i___iiikiiiii/ ii 'iI I!!i !il

"" I' ": "':::I :: : ..... /-za::: : :7::: : ;';'] ' ' ' I ' ' ' ] " '"

-14 -10 -6 -2 2 6 1 0

AOA (deg)

0.000

-0.002

"" -0.004-

--0.006

............... I ' ' I'''1' ' '

iil :i:! :[ i:!i: I:I[Z ]i : :il/ ..... --Y-_°

ii_!!i-i_/t i__ ii!:\? ::: ::i:":"

-0.008 ..........-18 -14 -10 -6 -2 2 6 10

AOA (deg)

Figure 22 Elevator Hinge Moment Effectiveness (V=100 kts, Surface Taps)

NASA/CR--2000-209921/VOL l 27

d

C-

q_

0

0

o_

DHC-6 Tailplane OSU 7X1 0 Wind Tunnel DataBaseline Surface Taps

1.6 .... , ! ! , ! !

------0-- re= 0.0, V=60 kts ;I" _ :

1.2 + #e= o.o, v=l oo kts ' [ ' _'' ; _"---E--- _e=-zo.o,v=eokts : ; : // I----:--i--_... _ ! !---_-- .=-_o.o,v.,oo_. i .... i ....=;¢ ...... ........ -_'

0.8 -- -- _ -- _e= 14- 2, V=60 kts

_e- 14.2, V--lO0 kts_ .:. .

o.o--i........................5-0.4

-0.8 i _r_'-

-1.2 I__: _

-1.6

-2.0-24 -20 -1 6 -12 -8 -4 0 4 8 12 1 6 20

Angle-of-Attac k (deg)

Figure 23 Velocity Effects, Baseline, Lift Characteristics

J

c--

O

O

DHC--6 Tailplane OSU 7X1 0 Wind Tunnel DataLEWlCE Ice Shape Surface Taps

-1 .6

-1 .2

-0.8

-0.4

0.0

0.4

0.8

i_212

Figure 24

8 4 0 -4 -8 -12 -1 6 -20

Angle-of--Attack (deg)

Velocity Effects, LEWICE Ice Shape, Lift Characteristics

NASA/CR--2000-209921 ]VOL 1 28

z

_5

(L:,

C:

CL_

EO

v

DHC-5 TailploneB,sseline

':OSU 7X1 O WFnd Tunnel =,,_'s;s,L

Su rfac _ To,._,_'-,_

Figure 25 Velocity Effects, Baseline, Hinge Characteristics

DHC-6Toilslone ,S,S_I 7 iL?,"A/Tnd--unnel E:'ot,s

LEWlCE ice Shape Surf,sceTaps

0.2_

z

C'.!@

()L-

O

o 0.0

EC'

z

¢' -0.I

I

-L;'.2-20 -1 6 -12 - 8 - 4 C,

Ang e-of-A+tack(,deg)

i4 F, 1 _

Figure 26 Velocity Effects, LEWICE Ice Shape, Hinge Moment Characteristics

NASA/CR--2000-209921/VOL 1 29

DHC-6 Tailplane OSU 7X1 0 Wind Tunnel DataBaseline V=I O0 kts

• , :'l'.....................- ;#- _i I ....:,1 2 _ _ Oe- 0,0. Belt Taps I• - - • - - ¢_e-- - 20.0, Surface Tooa

--o--,,--2o.0,Be,<,o0,..........L// ............o,8 -.-- _.-,,._,_o_oo.,o., i ' ' i)_ -'-_- --_-- _.=_4.2,B=.Top= i : / _ -"_m i........ E..... J • _ "'r " : .... r "_ d" ..... : ....... : •

: i : i : / i

0.4 f : : :c _i ...... i . i....................... i ..........

0.0 i i a

_- i i ij

o_ -o.4...:....i..._,;-0.8

-_.6 " =_::Sl j:--P:,....i I

" I ....... !..........i 2 I 0

I

-24 -20 -16 -12 -8 -4 0 4 8 12 16 20

Angle-of-Attac k (deg)

Figure 27 Tap Line Comparison, Baseline, Lift Characteristics

d

C

@

O

,+_@

O

LD

M.-

_J

DHC-6 Tailplane OSU 7X1 0 Wind Tunnel DataLEWICE Ice Shape V=I O0 kts

I , 2 __L__J___'__L__

-. "_ 6¢-O.O. SurfaccTaps

_1¢= 0.0, Belt Taps

0.8 _ --I-- 8e--20.O. SurfaceTaos- - O - " _e- -20.0, Belt T_0s

--k--- 6I- 14.2, Surface Topm

--L_-- - _e= 14.2, Belt Tap=

o.4 .....................

0.o i :,

-0.4 ............... _

• : .... ; ........ i=_4

-1.2

-1.6-20

! ! _ ! , !

, .._._,

! :-_>

-16 -12 -8 -4 0 4 8 12

Angle-of-Attac k (deg)

Figure 28 Tap Line Comparison, LEWICE Ice Shape, Lift Characteristics

NASA/CR--2000-209921/VOL 1 30

}HC 6Tailp!ane OSU 7XlOWindTunnel Dat,s_aselTne V=IO0 kts

0._

I

o 0.2

©

_, 0._Q,

0

©

E @@©

©

c -01

' [-0.2 ; ! i i ,-24 -20 -16 -12 -8 4 0 4 8 12 16 20

Angle-of-AttacK (deg)

Figure 29 Tap Line Comparison, Baseline, Hinge Moment Characteristics

Figure 30

DHC-6 Tailplane OSU 7X1 0 Wind Tunnel DataLEWlCE Ice Shape V=I O0 kts

T

CP

C

©CP

f-

8o

@

c-

-I-

0.2

0.1

0.0

_-........... ' ...... ' ........... ' .......... ' ......... ' " ----0-- 6e- 0.0, Surface Taps

r ..... ......... i ........ '--." -, .... i .... --t .... i ...... + ,.-0.0, ..It Tap.

.... : ...... i .......... u:.:.-_l .... .........I!j : ....... --l-- _,--20.0.so_oo,Top,....... , ..................... --,I_'- - _e- 14.2, SuOace Tap_

; : "_ : --_--- _e= 1 4.2, Bert Tapsi _ m

.. : ........ ; ........... i ......... i ..... -L:_'......... i .......... i.......

_ 0 _ 1

t

.... i ........ i ........... i ...... i .......... i .......... i .......... _....... : ....

-0.2 : _ _ _ i i _ i-20 -1 6 -1 2 -8 -4 0 ¢ 8

Angle-of-Attcc k (deg)

2

Tap Line Comparison, LEWICE Ice Shape, Hinge Moment Characteristics

NASA/CR--2000-209921 ]VOL 1 31

1,60

1.15

0.70

0.25

-0.20

-0.65

-1.10

-1.55

-2.00

" - • .... 6e=-26.6 Repeot

6e= D.D ', ..... ', - - .,_- S- '---_ ..... 6_=0.0Rep_o_ ' I / ' ' i,::_'''_ ,,

6e-- '14.2_ , -

!7 '----,--- _o=_4.2R_p_t J:. :i_,,,:._"' '• . • i , r : . . ,

0.3.o .= ..... ,.. ...... ,....... ,....... , ....... , ....... , ....

............ _............................................ •- - 6e--2e_6 n_t

.. - :- ..... '- -_ . . :....... ;....... ',....... ;....... : ........ 4,- - _o--:,e_e_,,:tOe= 0.0

0.2 l,.: ....... ! i I i ! .......... ........_,-oo,,..," " "' ............ ' ...... '" _ _l,=14.2

" " ._" r " : "" _-m:--_l- - : ....... '. "I ": " I : I --,i,---- _-14.2 ,,_,,._..... p,F ..... I '_-"PF" - _-i ...... '...... I ...... I"

...':_/4...':......:.._L..:..!'.__._.=...:...I...:._J......_...................0.1 , .... 1_,,,: .,, , ....

0.0...: ..... :.._.. : ..=..,.,....._,._- "'.-w.... : ...: ...... :..... T'-:2..._:%.'...... :..... ; .... __.

': .... f ..... :...... :...... :..... : , ,'x! ........:,-0.1 ..? ..... ?...... !....... !....... !....... !....... !...... !....... !..x@.i ..... i

!!iii[i!i!!! i!fiiiliiiii !!iiii i!iiiiiiiii!!i-24 -20 -16 --12 -8 -4- 0 4 8 12 16 20

Figure 31 Baseline Repeatability, V=100 kts, Surface Taps

NASAJCR--2000-209921/VOL 1 32

1.60

1.15

0.7O

0.25

-0.20

-0.65

-1.1 0

-1.55

-2.00

--r4-- 6e--26,6 ' '- - i- - 6e--26.B Repeet .....- - • - - 6e=-26.5 Repeat

........_-_........ 6e=O0 ;- ,-.'..-r -:- -,---i .... --'---,---.---,.--:--

+ Be= 0.0 Ropea_ _',--J _ -.-4i j t , , : . .

L_IT/.... i....... f" : " " " ....... ;- - i" -:..... i .... t ..... : ..... ;....

.: ..... :...... :....... :....... :..:,::- ..... : ..... :, 4 d./

l

/ ....

.t -i-I'.!:.,i .ti" : : : :

i i r _ i ' I

0.5 ....... : [ I I ]

---:- ..... :.... --: ....... : , : ",- ":-- ,- :- " _, --El-- _--2e.8...... L ..... i....... i....... i ....... l ....... i ...... -_-i -- III- - _._,, - 26.8 l_il_t

... :- . - , . :... ,.. - :....... : ....... :....... : ....... : ..... _I' .... _o--26-6 Rep_ot

0.2 ..... ;.... .-oo":I7I ", _-_,I"I"I ----v--- _-14.2

:.:---I: :l:: ::::I::::::::::::::::::::::::::::--'-.:___._:...i... i--_L..!...I..... .':1_,.,,,...I..... J.-:- .,.-:.-_-. _.... ,-0.1 " ' ' ' -'" .....

:::i:: ::i:: ::::::::::::::::::::::::::::::::::::::::::::::::::::::

o.o :..i i. :.... i..-_i:_:....:-., k ..... i.... i .... i..

-0.1 ........

...: ...... :...... :....... :....... :....... :....... : ...... :....... . ........... ;..-0.2 ......

--24 -20 -16 --12 -8 -4- 0 4 8 12 16 20

Figure 32 LEWICE Ice Shape Repeatability, V=100 kts, Surface Taps

NASA/CR--2000-209921/VOL 1 33

',_..... __, _ ......._;,J: 7 Y, 1 u,_......4'_.... 7 u _.... ,,c ! _; a : <_

1.2

Argie- of-At t':;c k _,,,_ ,9 ,_

Figure 33 Section 2 Lift Characteristics Comparison

< >

(i)

0

Cil

0

©

i_tHO -- EbToiis one 0<' " ...... '_ I.; ©c.tc_, _t,J ,"Vj U ';Vl _ _@

;).f

0.0

I- ...........................

- - = J .... ] ' ...i -t--- oe= 20 ?._'_,;:.-,r.-I--- _e= 0C _:._el:r!e

---A--- 4e= 4 Z. _-3o-:'elm_

.... v%4.

24 -20 -'6 --1_! -8 --4 ,L- 4

Figure 34 Section 2 Hinge Moment Comparison

8 12

NASA/CR--2000-209921 ]VOL 1 34

Figure 35 Baseline Tuft Flow Visualization, o_ = -8.0 °, _e " 0"0°

NASAJCR--2000-209921/VOL 1 35

Figure 36 S&C Ice Shape Tuft Flow Visualization, o_ =-8.0°,$e=0.0 °

NASA/CR--2000-209921/VOL 1 36

C_

rd

O

(.3

¢]

CO

L_

£L

c_LP

-4

-3

-2

1

DbC-6 Tailplane OSU 7X1 0 Wind Tunnel DataRur" 62 V=I O0 kts, Surface Taps

+

C2

2-01 0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9

x/c

Figure 37 Baseline Cp Distribution, ec =-8.0 °, _e----0.0 °

:.0

DHC-6Tailpane OSU7XlOWindTunnel DateRun 60, V=100 kts, Surface Taps

-4

3"d

._u 2©

0

(.3

©

o1,

G_

¢£(.3

0 Stob, S_c1

• Stab• Pres.

+ Boot To _ --

r " " _ ] 0 2 E4ev.Suct ', , , . . , i _ , • Elev Pres

.... i 1. . _FI . .

_9o_)O0o ; :'. .... '..... n _ ...... - , , I

0

p,,,,,.,"(' " I '

• [ ,

l

,0,1 0.O 0.1 0.2 0.3 0.4- 0.5 0.6 0.7 0.8 0.9 1.0

x/c

Figure 38 S&C Ice Shape Cp Distribution, e_ =-8.0 °, _}e=0.0 °

NASA]CR--2000-209921 ]VOL 1 37

DHC-6TaTipiane OSU 7×lC, W[ndTunr, e Date5 Hole Probe q indie,ation,\/=dO0 kts, 8e,-O.O

8 _ 2 1 6 2 0

Figure 39 5-Hole Probe Dynamic Pressure Comparison

DHC-6 Tailplsne OSU 7XI 0 W]rid TL. nrsei Dctc5-Hoie Probe AOA indication, V=100 _,ts, 5e=0.O

{/1

t._

C,"l

©

cl

c_

0.2

0,1

0.0

' I ' ! '

I , ,i j,, :

-0.t

-0.2

--0,5-24 -20 -

-i ......

6 -1 2 -8 -4 0 4 #3 1 2 1 6 20

Angle-of-Attack (,deg)

Figure 40 5-Hole Probe Pressure Differential due to Angle-of Attack

NASAJCR--2000-209921/VOL 1 38

DHC-6 Tailplane OSU 7X1 0 Wind Tunnel Data5-Hole Probe Sideslip Indication, V=I O0 kts, 6e=O.O

Q_

v

@I,_

03

O3

@k_

EL

0

©

d_

0.04

0.02

0.00

I ! ! ! ! I

• ' --O'_ Baseline i "

LEWlCEi

]

Angle-of-Attac k (deg)

Figure 41 5-Hole Probe Pressure Differential due to Sideslip

NASA/CR--2000-209921/VOL 1 39

8. CONCLUSIONS

A set of data quantifying the 2-D aerodynamics of the DHC-6 Twin Otter tailplane,

uniced, and with representative ice shapes, has been obtained. Instrumentation that will be used

during future flight testing has been verified to be effective. The results obtained are consistent

with expected aerodynamic trends and observed responses during typical pushover maneuvers.

The effects of icing that would have the greatest effect on aircraft stability and control

were shown in the changes in lift and hinge moment. For the uniced airfoil with Be=0 ° elevator

deflection, CLmax=-l.3 and AOAst_H=-18.6 °. With the LEWICE ice shape, these values were

reduced to CLmax=-0.64 and AOAst_ll=-8.2 °. Hinge moment data for the LEWICE ice shape at

6e=-20 ° showed that CH increased by ACH=0.02 (trailing edge down) about AAOA=2.3 ° before

the stall. These results supported observed responses during typical pushover maneuvers.

Reynolds number variation did not result in any strong characteristic differences in the

coefficient data with either ice shape installed. Reynolds number effects were seen for the uniced

baseline airfoil, with a lowering of stall AOA by approximately AAOA=2 ° and a decrease in liftcoefficient at the slower 60 kts velocities.

The elevator extension and large slot of the baseline airfoil section had a significant effect

on the pressure distributions. This was verified by testing of the secondary airfoil section, which

showed increased lift, CLma×= 1.35, and a decreased negative angle of attack at stall, AOA=- 16.5 °.

Hinge moment also increased, approximately ACH=0.04 for 5_---0.0 ° and -20.0 ° near stall. All of

the components that will be used in future flight testing showed acceptable fidelity. The pressure

belt should be a good source of data, however, care must be taken when evaluating these results,

as coefficient values can be lowered if the belt is immersed in largely separated flow. The 5-Hole

Probe data showed the expected trends with velocity, angle of attack, and sideslip. Differences in

dynamic pressures were seen between the tunnel and the probe values, as well as between the

baseline and LEWICE configuration data. Calibration of the probe, at a later date, will allow

more complete assessment of the probe accuracy and configuration dependencies.

These results have given an effective set of data with which to begin an analysis of

tailplane aerodynamics in icing conditions. The data will allow the tailplane aerodynamics during

simulated maneuvering flight to be quantified. The flight test instrumentation will allow

verification and refinement of these aerodynamics during flight test maneuvers. This combination

of simulation and flight test will allow a more complete understanding of the contribution of the

tailplane to the aircrafts motion during maneuvering flight. This understanding will be needed,

and fully utilized, in the development of maneuvers to identify aircraft susceptible to tailplane

stall in icing conditions.

NASA/CR--2000-209921/VOL 1 40

Appendix A

Tap Locations

Section 1, Clean

TAP# X/C TAP# X/C

1 -.0011 31 .3000

2 .0003 32 .3500

3 .0001 33 .3500

4 .0041 34 .4000

5 .0039 35 .4000

6 .0100 36 .4500

7 .0100 37 .4500

8 .0175 38 .5000

9 .0169 39 .5000

10 .0250 40 .5226

11 .0250 41 .5243

12 .0373 42 .5180

13 .0380 43 .5185

14 .0500 52 .5261

15 .0500 53 .5363

16 .0750 54 .5354

17 .0750 55 .5462

18 .1000 56 .5449

19 .1000 57 .5759

20 .1256 58 .5741

21 .1260 59 .5955

22 .1500 60 .5936

23 .1500 61 .6152

24 .1753 62 .6131

25 .1750 63 .6346

26 .2000 64 .6326

27 .2000 70 .6569

28 .2500 71 .6589

29 .2500 72 .7009

30 .3000 73 .7026

TAP # X/C

74 .7448

75 .7463

76 .7888

77 .7901

78 .8327

79 .8338

80 .8767

81 .8775

82 .9207

83 .9212

84 .9646

85 .9650

86 1.0000

Note s:

1) Stabilizer taps are #1-43, elevator taps are #52-86

2) Surface Taps do not include #63, and #64.

3) Belt Taps do not include #42, #43, #57, #63, #74, #78, and #82.

NASA/CR--2000-20992 I/VOL 1 41

Section 2, Clean

TAP# X/C TAP# X/C

1 -.0011 31 .3000

2 .0003 32 .3500

3 .0001 33 .3500

4 .0041 34 .4000

5 .0039 35 .4000

6 .0100 36 .4500

7 .0100 37 .4500

8 .0175 38 .5000

9 .0169 39 .5000

10 .0250 44 .5500

11 .0250 45 .5500

12 .0373 46 .6000

13 .0380 47 .6000

14 .0500 48 .6173

15 .0500 49 .6086

16 .0750 50 .6084

17 .0750 51 .6089

18 .1000 65 .6119

19 .1000 66 .6164

20 .1256 67 .6174

21 .1260 68 .6297

22 .1500 69 .6314

23 .1500 70 .6569

24 .1753 71 .6589

25 .1750 72 .7009

26 .2000 73 .7026

27 .2000 74 .7448

28 .2500 75 .7463

29 .2500 76 .7888

30 .3000 77 .7901

TAP # X/C

78 .8327

79 .8338

80 .8767

81 .8775

82 .9207

83 .9212

84 .9646

85 .9650

86 1.0000

Notes:

1) Stabilizer taps are #1-51, elevator taps are #65-86

2) Surface Taps do not include #76.

3) Belt Taps do not include #63 through #86.

NAS,_dCR--2000-209921/VOL 1 42

S&CIce ShapeTaps

TAP# X/C Y/C

1 -.0084 .00302 -.0027 .02013 -.0014 -.01504 -.0018 .03155 -.0019 -.02646 .0088 .03087 .0088 -.02758 .0175 .02869 .0175 -.0258

Notes:1)TheseTapsreplaceSurfaceTaps#1 through#I 1.2) TheseTapsreplaceBeltTaps#1 through#9.

LEWICE IceShapeTaps

TAP # X/C Y/C

1 -.0104 .00292 -.0207 .03383 -.0165 -.02824 -.0044 .02395 -.0044 -.0213

Notes:1)TheseTapsreplaceSurfaceTaps#1through#5,and#7.2)TheseTapsreplaceBeltTaps#1through#5.

NASA/CR--2000-209921]VOL 1 43

Y/C

0.1

0.05

0

-0,05

-01

-0,1

Section 1 (Baseline)

I

10 'Tlk I _I0_

11/ 19 0 0

i

0 0.1 0.2 0.3

X/C

X 64 76

41. 0_-: 0 0

o ,,_'-' / 7-_\ I_1 57 e_ 71 1'7

[£3

o o/": j

m/

0.4 0.5 0.6 0.7 0.8 0.9 I

01

0.05

Y/C 0

-005

-0.1

-0.1

Section 2

G

0_:_ 0

0 0.1 0.2 0.3

'I, " ,70 Io o _ '. _ ®°k°L 6o,

, 51- _ • 0 0

o , o, ,oj ,,0.4 0.5 0.6 07 0.8

X/C

DHC-6 Tailplane Airfoil Section Tap Locations

09 1

NASA/CR--2000-20992 I/VOL 1 44

>-

-1

-2

LEWICEIceShape

/\

JJ

-3.2 .1 0 1 2 3

_c{%}

00

-I

-2

.31-2

S&Ck:eS_e

• " " " ' .... i .............. " " " " "

5<

6

f/

_J1

J,*,[,l*, ll,,

2 3 4

Ice Shape Tap Locations

NAS A/CR--2000-209921/VOL 1 45

Start of Test 09-01-94, End of

Run Configuration

01 Baseline

02

03

04

O5

06

07

O8

09

l0

ll

12

13

14

15

16

17

18

19

20

21

22

23

24

25

26

27

28

29

30

31

32

33

34

35

36

Appendix B

Run Log

Test 09-20-94

Velocity Elevator AOA /Remarks

(kts) (deg) (deg)

60.0 0.0

" 100.0

" 60.0

" 100.0

" 60.0

" 100.0

Baseline 60.0

" 100.0

" 60.0

" 100.0

" 100.0

" 60.0

" 100.0

" 60.0

" 100.0

S&C Ice Shape 60.0

" 100.0

" 60.0

" 100.0

" 60.0

" 100.0

" 60.0

" 100.0

" 100.0

" 60.0

" 100.0

" 60.0

" 100.0

LEWICE Ice Shape 60.0

" 100.0

" 60.0

" 100.0

0.0

-10.0tl

-20.0If

-20.0

-26.6II

-26.6

10.0

14.2

14.2

_t

10.0

0.0ii

-10.0

-20.011

-26.6ft

-26.6

-20.0

18,16,14,12,10,8,6,4,2,0,-2,-4,-6,-8,

-10,-12,-14,-16,-18

12,8,4,0,-4,-8,-12,-14,-16,-18

- 18,-20,-22

12,8,4,0,-4,-8,- 12,- 14,- 16,- 18

Check Run

Check Run

12,8,4,0,-4,-8,-12,-14,-16,-18

12,8,4,0,-4,-8,- 12,- 14,- 16

12,8,4,0,-4

-8,-12,-14,-16

12,8,4,0,-4,-8,- 12,- 14,- 16,- 18

II

12,8A,0,-4,-8,- 10,- 12,- 14,

-16,-18

8,4,0,-4,-8,-10,-12,-14,-16

8,4,0,-4,-8,- 12,- 14,- 16

8,4,0,-4,-6,-8,- 10,- 12,- 14wt

vv

Bad run

8,4,0,-4,-6,-8,- 10,- 12,- 14

8,4,0,-2,-4,-6,-8,- 10,- 12Tt

T*

*v

t*

tt

Bad run

NASA/CR--2000-209921 ]VOL 1 47

Run Configuration Velocity Elevator

(kts) (deg)

AOA /Remarks

(deg)

37 LEWICE Ice Shape 60.0 - 10.0

38 " 100.0 "

39

40 " 60.0 0.0

41 " 100.0 "

42 " 60.0 10.0

43 " 100.0 "

44 " 60.0 14.2

45 " 100.0 "

46 ......

47 " 100.0 0.0

48 .... -26.6

49 ......

50 Baseline ....

51 ......

52 .... -10.0

53 " 60.0 "

54 " 60.0 0.0

55 " 100.0 "

56 " 100.0 "

57 " 100.0 14.2

58 S&C Ice Shape 60.0 "

59 " 100.0 "

60 .... 0.0_ _v In

61 .... -26.6

Baseline 60.0 "

" 60.0 0.0

62 Baseline, NP 100.0 "

63 Section 2, NP 60.0 "

64 .... -20.0

65 .... 14.2

66 .... 0.0

8,4,0,-2,-4,-6,-8,- 10,- 12

Not Used

8,6,4,2,0,-2,-4,-6,-8,- 10,

-12,-14w

8,4,0,-4,-8,- 10,- 12,- 14,- 16*t

8,4,0,-4,-8,- 10,- 12,- 14,- 16

Repeat of 045

Repeat of 041

Repeat of 033

Repeat 2 of 033

Repeat of 013

Repeat 2 of 013

Repeat of 004

Repeat of 003

Repeat of 001

Repeat of 002

Repeat of 055

Repeat of 018

Repeat of 019

Repeat of 020

Repeat of 024Flow Viz

Repeat of 031, Flow VizFlow Viz

Flow Viz

16,12,8,4,0,-4,-8,- 12,- 14,- 16,

- 18,-20, Flow Viz

8,4,0,-4,-8,- 10,- 12,- 14,- 16,Flow Viztu

ii

- 16,- 18,-20,-22,Flow Viz

Notes:

1. Runs 01 through 07 had an inoperative scanivalve.2. Run 19 had the 5-Hole Probe cover installed.

3. Wake probe sweep runs - 0001,0002,0003,0023,0024,0040,0041.00634. Wall Statics available for runs 0035 and up, only.5. Flow Viz indicates more complete tufting and documentation.6. Post run AOA=0.0 ° point for all runs after 009 not indicated.7. NP = No 5-Hole Probe.

NASA/CR--2000-209921/VOL 1 48

Appendix C

Video Log

This log contains a listing of the approximate time for the video taken during testing. To

conserve tape, only approximately 15 seconds of each data point were recorded. The times listed

below are the approximate start times for each run.

Run 0001 Run 0002 Run 0003/5

Date 09/08/94 Date 09/08/94 Date 09/09/94

AOA Time AOA Time AOA Time

18 1324 16 1629 12 0931

16 1328 12 1634 8 0937

14 1334 8 1639 4 0942

12 1341 4 1644 0 0947

10 1346 0 1649 -4 0952

8 1352 -4 1655 -8 0958

6 1357 -8 1700 -12 1004

4 1402 -10 1707 -14 1009

2 1407 -12 1710 -16 1013

0 1412 -14 1715 -18 1018

-2 1416 -16 1721 -18 1153

-4 1422 -20 1156

-6 1426 -22 1200

-8 1431

-10 1435

-12 1442

-14 1446

-16 1452

-18 1456

Run 0004 Run 0006 Run 0007

Date 09/09/94 Date 09/08/94 Date 09/09/94

AOA Time AOA Time AOA Time

12 1103 12 1242 12 1328

8 1106 8 1246 8 1334

4 1113 4 1249 4 1335

0 1114 0 1253 0 1339

-4 1117 -4 1256 -4 1342

-8 1120 -8 1259 -8 1346

-12 1123 -12 1303 -12 1350

-14 1127 -14 1306 -14 1354

-16 1130 -16 1309 -16 1356

-18 1133 -18 1313 -18 1359

-20 1316 -20 1403

NASA/CR--2000-209921/VOL 1 49

Run 0010Date09/12/94AOA Time12 15538 15574 16000 1605-4 1606-8 1610

-12 1613-14 1616-16 1620-18 1623

Run 0011Date09/12/94AOA Time12 16398 16434 16470 1650-4 1653-8 1656-12 1700-14 1704-16 1707-18 1711

Run 0012Date09/12/94AOA Time12 17588 18014 18050 1808-4 1812-8 1815-12 1818-14 1182-16 1825

Run 0013/14Date09/12/13/94AOA Time12 18368 18404 18430 1847-4 1850-8 0929-12 0932-14 0936-16 0939

Run 0015Date09/13/94AOA Time12 10158 10194 10220 1026-4 1029-8 1032

-12 1035-14 1038-16 1042-18 1745

Run 0016Date09/13/94AOA Time12 ll018 11054 11080 1111-4 1115-8 1118-12 1122-14 1125-16 1128-18 1045

Run 0017Date09/12/13/94AOA Time12 12208 12244 12280 1232-4 1235-8 1238-12 1241-14 1244-16 1247-18 1251

Run 0018Date09/13/94AOA Time12 13058 13094 13120 1316-4 1319-8 1322-12 1326-14 1329-16 1332-18 1336

Run 0019Date09/14/94AOA Time12 10068 10104 10130 1017-4 1020-8 1024-12 1027-14 1030-16 1037-18 1040

NASAJCR--2000-209921/VOL 1 50

Run 0020

Date 09/14/94

AOA Time

8 1110

4 1114

0 1118

-4 1121

-8 1124

-10 1127

-12 1131

-14 1135

-16 1138

Run 0021

Date 09/14/94

AOA Time

8 1216

4 1220

0 1223

-4 1227

-8 1230

-10 1233

-12 1236

-16 1240

-18 1243

Run 0022

Date 09/14/94

AOA Time

8 1256

4 1300

0 1304

-4 1307

-8 1310

-10 1313

-12 1317

-16 1320

-18 1323

Run 0023

Date 09/14/94

AOA Time

8 1400

4 1406

0 1414

-4 1415

-6 1421

-8 1428

-10 1430

-12 1435

-14 1441

Run 0024

Dme 09/14/94

AOA Time

8 1458

4 1505

0 1509

-4 1515

-6 1520

-8 1525

-10 1530

-12 1537

-14 1541

Run 0025

Date 09/14/94

AOA Time

8 1628

4 1633

0 1635

-4 1639

-6 1642

-8 1645

-10 1649

-12 1652

-14 1655

Run 0026/27

Date 09/14/94

AOA Time

8 1708

4 1711

0 1715

-4 1718

-6 1722

-8 1725

-10 1728

-12 1731

-14 1849

Run 0028

Date 09/14/94

AOA Time

8 1929

4 1932

0 1938

-2 1939

-4 1943

-6 1946

-8 1949

-10 1953

-12 1956

Run 0029

Date 09/14/94

AOA Time

8 2010

4 2013

0 2016

-2 2020

-4 2023

-6 2026

-8 2029

-10 2033

-12 2036

NASAJCR---2000-20992 I/VOL 1 51

Run 0030

Date 09/15/94

AOA Time

8 0910

4 0913

0 0918

-2 0921

-4 0923

-6 0927

-8 0930

-10 0934

-12 0937

Run 0031

Date 09/15/94

AOA Time

8 0955

4 1000

0 1005

-2 1007

-4 1010

-6 1013

-8 1017

-10 1021

-12 1023

Run 0032

Date 09/15/94

AOA Time

8 1455

4 1459

0 1503

-2 1506

-4 1510

-6 1513

-8 1516

-10 1519

-12 1522

Run 0033

Date 09/15/94

AOA Time

8 1537

4 1541

0 1545

-2 1548

-4 1551

-6 1554

-8 1558

-10 1601

-12 1605

Run 0034

Date 09/15/94

AOA Time

8 1658

4 1702

0 1705

-2 1708

-4 1711

-6 1715

-8 1718

-10 1722

-12 1725

Run 0035

Date 09/15/94

AOA Time

8 1832

4 1836

0 1839

-2 1843

-4 1846

-6 1850

-8 1853

-10 1856

Run 0037

Dme 09/15/94

AOA Time

8 2012

4 2015

0 2018

-2 2021

-4 2025

-6 2028

-8 2031

-10 2034

-12 2037

Run 0038

Date 09/15/94

AOA Time

8 2053

4 2056

0 2100

-2 2103

-4 2107

-6 21t0

-8 2114

-10 2117

-12 2120

Run 0040

Date 09/16/94

AOA Time

8 0904

6 0910

4 0915

2 0920

0 0925

-2 0930

-4 0934

-6 0940

-8 0945

-10 0950

-12 0955

-14 1000

NASA/CR--2000-209921]VOLI 52

Run 0041Date09/16/94AOA Time8 11596 12054 12102 12150 1220-2 1225-4 1230-6 1235-8 1241-10 1245-12 1250-14 1256

Run 0042Date09/15/94AOA Time8 14004 14050 1409-4 1412-8 1415-10 1418-12 1421-14 1425-16 1428

Run 0043Date09/16/94AOA Time

8 16104 16140 1617-4 1620-8 1624-10 1627-12 1630-14 1633-16 1637

Run 0044Date09/16/94AOA Time8 17304 17350 1738-4 1741-8 1745-10 1748-12 1752-14 1755-16 1758

Run 0045Date09/15/94AOA Time

8 18284 18330 1835-4 1838-8 1842-10 1845-12 1848-14 1852-16 1855

NASA/CR--2000-20992I/VOL1 53

Appendix DSolid Wall Correction Calculations

The effect of the walls on the 2-dimensional airfoil data must be considered. Corrections

have been applied to the raw wind tunnel data using the methods of Ref. D 1. For this test, three

phenomena have been taken into account:

Solid Blockage

Solid blockage creates an increase of velocity due to the reduction of the area through

which the air must flow. To quantify this effect, computations have been made by considering

the two-dimensional model as a cylinder. This cylinder can be simulated as a doublet of strength

/a = 2rWa 2 (a : cylinder radius), with the walls represented by the streamlines created by matching

doublets, on each side of the cylinder. Therefore,

2AV a

V,, tl:(The subscript u indicates uncorrected data)

After doing a summation, we obtain,

with:

h : tunnel height

c : model chord

/l.2 C -

!

A=--_-16[.:Y[(1-P(l+dVllSdXcL-dx)l c

x, y: airfoil coordinates, P its no-camber,

symmetrical, pressure distribution.

Usually A is given by a graph

In our case, A = 0.22

NASA/CR--2000-209921/VOL 1 55

Wake Blockage

Wake blockage creates a velocity increment at the model because of a pressure gradient

due to higher velocity which keeps the flow around the model. In computations, the wake is

simulated by a line source and the walls by an infinite vertical row of source-sink combinations.This reduces to:

AV CE.b -- -- Ca.

V,, 4h

We can, then, correct the tunnel conditions by these equations:

V = V,, (1 + e ) with :

q = q,(l+2e ) e= e_,, + e,,,,

Re = Re,(l + e )

From the dynamic pressure effect and the wake gradient term, we get:

C,_ = Cd,(l -3 e_,-2 e,t, )

Streamline Curvature

The lift and moment about the quarter chord of an airfoil are too large at a given angle of

attack (which is also too large). This is due to the fact that the airfoil seems to have more camber

because of the floor and ceiling.

Calculations of this effect are made by assuming that the airfoil can be approximated by a

single vortex at its quarter-chord point. The floor and ceiling are represented by a vertical row of

vortices, extending to infinity and with alternating signs. The load on the airfoil can be

decomposed as a flat plate loading, computed as an angle of attack correction, and an elliptical

loading, which gives the lift, pitching-moment and hinge-moment corrections.

It can be shown that the upwash induced at the half chord by the two images is:

AO_ =

2

1 c

NASA/CR--2000-209921/VOL 1 56

We can consider that (c/4) 2 is small compared to h 2, so it gives 6 equal to 6 previously found,

and we have then:

57. o( )cz=a,,+ 2:,r C_,+4C,,,_,

C_ = C_,(1-<y-2e )

GC_C,/ = C,,,L,,(1 -2e)+--

4 4 4

Hinge Moment

To establish a correction for the hinge moment, we follow the same analysis process

discussed above, but we consider only the flap, or elevator, instead of the entire airfoil. We are

now at a three-quarter-chord point, so we lose accuracy by taking (3c/4)2<<h 2 in the Aot

expression.

Therefore, including this 3c/4 term results in 6'= 0.9 6.

So,

C} -t

C,, = Ch,,(1- 2e )+-4-CI,-with:

C_f : lift coefficient of the elevator

References

D 1. Rae, W. and Pope, A., Low Speed Wind Tunnel Testing, Wiley-Interscience, 1984.

NASA/CR--2000-209921/VOL 1 57

Appendix E

Data Reduction Output File Formats

Uncorrected Coefficients

File Name: AA500XX.DAT

Run: 00XX

Column Value

1

2

3

4

5

6

AOA (degrees)

Elevator Deflection (degrees)

Mach Number

Reynolds Number (millions)

Q (psi)

Velocity (kts.)

From Surface Taps

7 CL (total)

8 CD (pressure only)

9 CM (1/4c)

10 CL (Elevator only)

11 C. (Elevator)

From Belt Taps

12 CL

13 CD (pressure only)

14 CM (1/4c)

15 CL (Elevator only)

16 Ca (Elevator)

17 Run Number

Corrected Coefficients using Surface Taps

File Name: SC500XX.DAT

Run: 00XX

Column Value

AOA (degrees)

Elevator Deflection (degrees)Mach Number

Reynolds Number (millions)

NASA/CR--2000-20992 I/VOL 1 59

5 Q (psi)

6 Velocity (kts.)

7 CL (total)

8 CD (pressure only)

9 CM (1/4c)

10 CH (Elevator)

Corrected Coefficients using Belt Taps

File Name: BC500XX.DAT

Run: 00XX

Column Value

1

2

3

4

5

6

7

8

9

10

AOA (degrees)

Elevator Deflection (degrees)

Mach Number

Reynolds Number (millions)

Q (psi)

Velocity (kts.)

CL (total)

CD (pressure only)

CM (1/4C)

C. (Elevator)

Probe Pressures

File Name: PP500XX.DAT

Run: 00XX

Column Value

1

2

3

4

5

6

7

8

9

AOA (degrees)

Elevator Deflection (degrees)

Mach Number

Reynolds Number (millions)

Q (psi)

Velocity (kts.)

P Total (psi.)

P Static (psi.)

P AOA 1 (psi.) (for AOA Suction side )

NAS A/CR--2000-209921/VOL 1 60

10

11

12

P AOA 2 (psi.) (for AOA Pressure side )

P Beta 1 (psi.) (for Sideslip R/H side )

P Beta 2 (psi.) (for Sideslip L/H side )

Wall Static Pressures

(Available for Runs 0035 and up, only. )

File Name: WS500XX.DAT

Run: 00XX

Column Value

1

2

3

4

5

6

7

8

9

10

11

12

AOA (degrees)

Elevator Deflection (degrees)

Mach Number

Reynolds Number (millions)

Q (psi)

Velocity (kts.)

Ps West Forward (psi.)

Ps West Center (psi.)

Ps West Aft (psi.)

Ps East Forward (psi.)

Ps East Center (psi.)

Ps East Aft (psi.)

NASA/CR--2000-20992 l/VOL 1 61

Appendix F

DHC-6 Twin Otter Tailplane Coefficient Data

Figure F.01

Figure F.02

Figure F.03

Figure F.04

Figure F.05

Figure F.06

Figure F.07

Figure F.08

Figure F.09

Figure F.10

Figure F.11

Figure F.12

Figure F.13

Figure F.14

Figure F.15

Figure F.16

Figure F.17

Figure F.18

Figure F.19

Figure F.20

Figure F.21

Figure F.22

Figure F.23

Figure F.24

Figure F.25

Figure F.26

Figure F.27

Figure F.28

Figure F.29

Figure F.30

Figure F.31

Figure F.32

Figure F.33

Figure F.34

Figure F.35

Figure F.36

Figure F.37

Figure F.38

Figure F.39

Baseline, 100 kts .......................................................................................................... 64

Baseline, 60 kts ............................................................................................................ 65

Baseline, 100 kts, Belt Taps ....................................................................................... 66

Baseline, 60 kts, Belt Taps ......................................................................................... 67

Lewice Ice Shape 100 kts ........................................................................................... 68

Lewice Ice Shape 60 kts ............................................................................................. 69

Lewice Ice Shape, 100 kts, Belt Taps ........................................................................ 70

Lewice Ice Shape, 60 kts, Belt Taps .......................................................................... 71

S&C Ice Shape, 100 kts .............................................................................................. 72

S&C Ice Shape, 60 kts ................................................................................................. 73

S&C Ice Shape, 100 kts, Belt Taps ............................................................................ 74

S&C Ice Shape, 60 kts, Belt Taps .............................................................................. 75

Ice Shape Effects 100 kts, Positive Elevator ............................................................ 76

Ice Shape Effects 100 kts, Negative Elevator .......................................................... 77

Ice Shape Effects 60 kts, Positive Elevator .............................................................. 78

Ice Shape Effects 60 kts, Negative Elevator ............................................................ 79

Ice Shape Effects Belt Taps, 100 kts, Positive Elevator ......................................... 80

Ice Shape Effects Belt Taps, 100 kts, Negative Elevator ........................................ 81

Ice Shape Effects Belt Taps, 60 kts, Positive Elevator ........................................... 82

Ice Shape Effects Belt Taps, 60 kts, Negative Elevator .......................................... 83

Velocity Effects, Baseline ............................................................................................ 84

Velocity Effects, Baseline, Belt Taps ......................................................................... 85

Velocity Effects, Lewice Ice Shape ............................................................................ 86

Velocity Effects, Lewice Ice Shape, Belt Taps ......................................................... 87

Velocity Effects, S&C Ice Shape ................................................................................ 88

Velocity Effects, S&C Ice Shape, Belt Taps ............................................................. 89

Tap Line Effects Baseline, 100 kts ............................................................................ 90

Tap Line Effects Baseline, 60 kts .............................................................................. 91

Tap Line Effects Lewice Ice Shape, 100 kts ............................................................ 92

Tap Line Effects Lewice Ice Shape, 60 kts .............................................................. 93

Tap Line Effects S&C Ice Shape, 100 kts ................................................................ 94

Tap Line Effects S&C Ice Shape, 60 kts .................................................................. 95

Repeatability, Baseline ................................................................................................ 96

Repeatability, Baseline, Belt Taps ............................................................................. 97

Repeatability, Lewice Ice Shape ................................................................................ 98

Repeatibility, Lewice Ice Shape, Belt Taps .............................................................. 99

Repeatability, S&C Ice Shape .................................................................................. 100

Repeatability, S&C Ice Shape, Belt Taps ............................................................... 101

Section 2 Comparison, V=100 kts, Surface Taps ................................................. 102

NASA/CR--2000-209921/VOL 1 63

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NASA/CR--2000-209921/VOL 1 102

Form ApprovedREPORT DOCUMENTATION PAGEOMB No. 0704-0188

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

September 2000 Final Contractor Report

4. TITLE AND SUBTITLE 5. FUNDING NUMBERS

DHC-6 Twin Otter Tailplane Airfoil Section Testing in the Ohio State University

7x10 Wind Tunnel

6. AUTHOR(S)

Dale Hiitner, Michael McKee, Karine La Nor, and Gerald Gregorek

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

Ohio State University Research Foundation

1960 Kenny Road

Columbus, Ohio 43210-1063

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

National Aeronautics and Space Administration

John H. Glenn Research Center at Lewis Field

Cleveland, Ohio 44135-3191

WU-548-21-23--00

NAG3-1574

8. PERFORMING ORGANIZATION

REPORT NUMBER

E-12159-1

10. SPONSORING/MONITORINGAGENCY REPORT NUMBER

NASA CR--2000-209921 -VOL1

11. SUPPLEMENTARY NOTES

Project Manager, Thomas Ratvasky, Turbomachinery and Propulsion Systems Division, NASA Glenn Research Center,

organization code 5840, (216) 433-3905.

12a. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified - Unlimited

Subject Categories: 02 and 05 Distribution: Nonstandard

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

12b. DISTRIBUTION CODE

13, ABSTRACT (Maximum 200 words)

Ice contaminated tailplane stall (ICTS) has been found to be responsible for 16 accidents with 139 fatalities over the last three decades, and is suspected

to have played a role in other accidents and incidents. The need for fundamental research in this area has been recognized at three international[

conferences sponsored by the FAA since 1991. In order to conduct such research, a joint NASA/FAA Tailplane Icing Program was formed in t994: the

Ohio State University has played an important role in this effort. The program employs icing tunnel testing, dry wind tunnel testing, flight test.,ng, and

analysis using a six-degrees-of-freedom computer code tailored to this problem. A central goal is to quantify the effect of tailplane icing on mrcraft

stability and control to aid in the analysis of flight test procedures to identify aircraft susceptibility to ICTS. This report contains the results ot testing of

a full scale 2D model of a tailplane section of NASA's Icing Research Aircraft, with and without ice shapes, in an Ohio State University 7×10 Low

Speed wind tunnel in 1994. The results have been integrated into a comprehensive database of aerodynamic coefficients and stability and control

derivatives that will permit detailed analysis of flight test results with the analytical computer program. The testing encompassed a full range of angles

of attack and elevator deflections, as well as two velocities to evaluate Reynolds number effects. Lift, drag, pitching moment, and hinge moment

coefficients were obtained. In addition, instrumentation for use during flight testing was verified to be effective, all components showing acceptable

fidelity. Comparison of clean and iced airfoil results show the ice shapes causing a significant decrease in the magnitude of CLmax (from -1.., to -0.64)

and associated stall angle (from -18.60 to -8.2°). Furthermore, the ice shapes caused an increase in hinge moment coefficient of approximatel v 0.02, the

change being markedly abrupt for one of the ice shapes. A noticeable effect of elevator deflection is that magnitude of the stall angle is decrea,;ed for

negative (upward) elevator deflections. All these result are consistent with observed tailplane phenomena, and constitute an effective set of data for

comprehensive analysis of ICTS.

14. SUBJECT TERMS

Aircraft icing; Tailplane icing; Airfoil performance

17. SECURITY CLASSIFICATIONOF REPORT

Unclassified

18. SECURITY CLASSIFICATIONOF THIS PAGE

Unclassified

19. SECURITY CLASSIRCATIONOF ABSTRACT

Unclassified

15. NUMBER OF PAGES

112

16. PRICE CODE

AQO20. LIMITATION OF ABSTRACT

NSN 7540-01-280-5500 Standard Form 298 (Rev. 2-89)

Prescribed by ANSI Std. Z39- ! 8298-102