DHC-6 Twin Otter Tailplane Airfoil Section Testing in the Ohio State University 7x10 Wind Tunnel

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

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 were taken 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, within 10 -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 were accomplished.

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.

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-20 -16 -12 -8 -4 0 4 8 12 16 20 AOA (deg) 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 sign conventions.

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 a negativeAOA below stall 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 baseline case. 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 measuredto

within 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

pressures starting at approximately80%chord. Overall, the tuff flow visualizationshowswell

defined2-D flow andgoodagreement with the Cpdistributions.

7.7 5-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.

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Deflection + Hinge Moment + Angle-of-Attack - Lift ( + Downforce) Figure 13 Axis Systems and Sign Conventions NASA/CR--2000-209921/VOL 1 22

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0.02 0.02 ....

":::1: ::1: ::1 :: :1: : :1: ::": :i] 0.00 : : : ' : : : ....... :':::' ;;';:i 0.00 ................

-18 -1¢ -10 -6 -2 2 5 10 -18 -14 -10 -6 -2 2 6 10 ^o_ ((_) AOA (deg) Figure 21 Elevator Lift Effectiveness (V= 100 kts, Surface Taps) 0000 0.0o0 .: " I " : ] : : ] :..:.; L: .: :.I .: : .:.J :.: .:.

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

ii ........... iiiiiililtliliill!! iiti!i_ili

_0 _002 -0002 i i ii i ii -: :-;-r-.-;-:--I-;-:-:-i ;-:-; I--: .:-:--I -: -. 1 ;-. -:- i,-_' -i-/i i i"-:, i! i _

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

-0004- ....

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

-0006 :.

_,:,_,_E%}% .... . : i :i i -, ! !i- .: : : .... 2.:..I.:.:..:.] :..: : L : 2.2 I : :-.J :-: : : : : : ' : : ;" ; , , -0.008 -0008 O00Q 0.000 ............... I ' ' I'''1' ' ' /'.!2 :.!.i ',.J L.!..].!!.!.IL!!.',.!L:.:.:.:.:..

iil :i:! :[ i:!i: I:I[Z ]i ::il/ ..... --Y-_° -0.002 -0 002 .:..:.:..Li.: ..: :.:.. ;.:..:/.:.: ;..:.: ;.--."::_-.:..

-0004 "" -0.004-

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

--0.006 -0 006 ii_ !!i-i_/t i__ i i!:\? ::: ::i:":" "" I' ": "':::I :: : ..... /-za ::: : :7::: : ;';'] ' ' ' I ' ' ' ] " '" -0.008 ..........

-0008 -18 -14 -10 -6 -2 2 6 10 -18 -14 -10 -6 -2 2 6 1 0 AOA (deg) AOA (deg) Figure 22 Elevator Hinge Moment Effectiveness (V=100 kts, Surface Taps) NASA/CR--2000-209921/VOL l 27 DHC-6 Tailplane OSU 7X1 0 Wind Tunnel Data Baseline 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 .... =;¢ ...... ........ -_' -- -- _ -- _e= 14- 2, V=60 kts 0.8 _e- 14.2, V--lO0 kts _ .:. .

d C- q_ -0.4 o_ -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 DHC--6 Tailplane OSU 7X1 0 Wind Tunnel Data LEWlCE Ice Shape Surface Taps -1 .6 -1 .2 -0.8 J c-- -0.4 0.0 O O 0.4 0.8 i_2 8 4 0 -4 -8 -12 -1 6 -20 Angle-of--Attack (deg) Figure 24 Velocity Effects, LEWICE Ice Shape, Lift Characteristics NASA/CR--2000-209921 ]VOL 1 28 ':OSU 7X1 O WFnd Tunnel =,,_'s;s, L DHC-5 Tailplone B,sseline Su rfac _ To,._,_'-,_ z _5 (L:, C: CL_ E O v 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 E C' z ¢' -0.I I i -L;'.2 4 F, 1 _ -20 -1 6 -12 - 8 - 4 C, Ang e-of-A+tack(,deg) 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 Data Baseline V=I O0 kts ....................- ;#- _i I .... :,

• , : 'l'. 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 i 2 I 0 -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 DHC-6 Tailplane OSU 7X1 0 Wind Tunnel Data LEWICE 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= d o.4 ..................... , .._._, C @ 0.o i :, O ,+_ @ O LD -0.4 ............... _ M.- _J ! :-_> • : .... ; ........ 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, © 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 DHC-6 Tailplane OSU 7X1 0 Wind Tunnel Data LEWlCE Ice Shape V=I O0 kts 0.2 _-........... ' ...... ' ........... ' .......... ' ......... ' " ----0-- 6e- 0.0, Surface Taps r ..... ......... i ........ '--." -, .... i .... --t .... i ...... + ,.-0.0, ..It Tap.

T CP ....... , ..................... --,I_'- - _e- 14.2, SuOace Tap_ .... :...... i .......... u:.:.-_l .... ......... I! j :....... --l-- _,--20.0.so_oo,Top, ; : "_ : --_--- _e= 1 4.2, Bert Taps i _ m 0.1 C .. : ........ ;........... i ......... i ..... -L:_' ......... i .......... i.......

-0.2 : _ _ _ i i _ i -20 -1 6 -1 2 -8 -4 0 ¢ 8 Angle-of-Attcc k (deg) Figure 30 Tap Line Comparison, LEWICE Ice Shape, Hinge Moment Characteristics NASA/CR--2000-209921 ]VOL 1 31 1,60 1.15 " - • .... 6e=-26.6 Repeot 6e= D.D ', ..... ', - - .,_- S- ' ---_ ..... 6_=0.0Rep_o_ ' I / ' ' i,::_'''_ ,, 0.70 6e-- '14.2_ , -

----,--- _o= _4.2R_p_t J:. :i_,,,:. _"' '

!7 '

• . • i , r : . . , 0.25 -0.20 -0.65 -1.10 -1.55 -2.00 0.3 .o .= ..... ,.. ...... ,....... ,....... , ....... , ....... , ....

............ _ ............................................ •- - 6e--2e_6 n_t .. - :- ..... '- -_ . . :....... ;....... ',....... ;....... : ........ 4,- - _o--:,e_e_,,:t Oe= 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!fi iiliiii i !!iiii i!i iii iiiii!!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 --r4-- 6e--26,6 ' ' - - i- - 6e--26.B Repeet .....

1.15 - - • - - 6e=-26.5 Repeat ........_-_........ 6e=O0 ;- ,-.'..-r -:- -,---i .... --'---,---.---,.--:-- + Be= 0.0 Ropea_ _',--J _ -.-4 0.7O i j t , , : . .

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

-0.20 4 d .: ..... : ...... : ....... : ....... :..:,::- ..... :..... :, ./ l -0.65 / ....

-1.1 0 .t -i-I'.!:.,i .ti" : : : : i i r _ i ' I -1.55 -2.00 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_tHO -- EbToiis one 0<' " ...... '_ I.; ©c.tc _, _t,J ,"Vj U ';Vl _ _@ I- ...........................

- - = J .... ] ' ...i -t--- oe= 20 ?._'_,;:.-,r.- I--- _e= 0C _:._el:r!e ---A--- 4e= 4 Z. _-3o-:'elm_ < > ;).f (i) Cil 0.0 © .... v%4.

24 -20 -'6 --1_! -8 --4 ,L- 4 8 12 Figure 34 Section 2 Hinge Moment Comparison 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 DbC-6 Tailplane OSU 7X1 0 Wind Tunnel Data Rur" 62 V=I O0 kts, Surface Taps -4 + -3 C2 C_ rd -2 O (.3 ¢] CO L_ £L c_ LP :.0 -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 ° DHC-6Tailpane OSU7XlOWindTunnel Date Run 60, V=100 kts, Surface Taps -4 0 Stob, S_c1 • Stab• Pres.

+ Boot To _ -- "d r " " _ ] 0 2 E4ev.Suct ' , , , . . , i _ , • Elev Pres ._u 2 .... i 1. . _FI . .

(.3 © .... n _ ...... - , , I o1, G_ p,,,,, .," ¢£ (' " I ' (.3 • [ , 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 Date 5 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 Dctc 5-Hoie Probe AOA indication, V=100 _,ts, 5e=0.O 0.2 ' I ' ! ' I , ,i j,, : 0,1 {/1 0.0 t._ C,"l © cl -0.t c_ -0.2 -i ......

--0,5 -24 -20 - 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 Data 5-Hole Probe Sideslip Indication, V=I O0 kts, 6e=O.O 0.04 I ! ! ! ! I • ' --O'_ Baseline i " i LEWlCE ] 0.02 Q_ v @ I,_ 0.00 O3 @ k_ EL © d_ 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 lift coefficient 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

Next: Appendix A →

Appendix A

Appendix A Tap Locations Section 1, Clean TAP # X/C TAP# X/C TAP# X/C 74 .7448 1 -.0011 31 .3000 75 .7463 2 .0003 32 .3500 3 .0001 33 .3500 76 .7888 4 .0041 34 .4000 77 .7901 78 .8327 5 .0039 35 .4000 79 .8338 6 .0100 36 .4500 80 .8767 7 .0100 37 .4500 8 .0175 38 .5000 81 .8775 9 .0169 39 .5000 82 .9207 10 .0250 40 .5226 83 .9212 84 .9646 11 .0250 41 .5243 85 .9650 12 .0373 42 .5180 13 .0380 43 .5185 86 1.0000 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 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

Next: Section 2, Clean →

Section 2, Clean

Section 2, Clean TAP# X/C TAP# X/C TAP # X/C 1 -.0011 31 .3000 78 .8327 2 .0003 32 .3500 79 .8338 80 .8767 3 .0001 33 .3500 4 .0041 34 .4000 81 .8775 5 .0039 35 .4000 82 .9207 6 .0100 36 .4500 83 .9212 7 .0100 37 .4500 84 .9646 8 .0175 38 .5000 85 .9650 9 .0169 39 .5000 86 1.0000 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 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&C Ice Shape Taps

TAP # X/C Y/C

1 -.0084 .0030

2 -.0027 .0201

3 -.0014 -.0150

4 -.0018 .0315

5 -.0019 -.0264

6 .0088 .0308

7 .0088 -.0275

8 .0175 .0286

9 .0175 -.0258

Notes: 1)TheseTapsreplaceSurfaceTaps#1 through#I 1.

2) TheseTapsreplaceBelt Taps#1 through#9.

LEWICE IceShape Taps

TAP # X/C Y/C

1 -.0104 .0029

2 -.0207 .0338

3 -.0165 -.0282

4 -.0044 .0239

5 -.0044 -.0213

Notes: 1)These Taps replace Surface Taps #1through #5,and #7.

2)These Taps replace BeltTaps #1through #5.

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

Next: Section 1 (Baseline) →

Section 1 (Baseline)

0.1 I X 64 76

[

10 'Tlk I _I0_ 0.05 £3 Y/C

o o/

41. 0_-: 0 0 ": j -0,05 m/

o ,,_'-' / 7-_\ I

11 / 19 0 0 _1 57 e_ 71 1'7 i -01 0.4 0.5 0.6 0.7 0.8 0.9 I -0,1 0 0.1 0.2 0.3 X/C

Section 2 'I, " ,70 I

0.05 G

o o _ '. _ ®°k°

L 6o, Y/C 0 , 51- _ • 0 0 -005 0_:_ 0 o , o, ,oj ,, -0.1 0.4 0.5 0.6 07 0.8 09 1 -0.1 0 0.1 0.2 0.3 X/C

DHC-6 Tailplane Airfoil Section Tap Locations

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

LEWICE Ice Shape

J J >-

/\

-1

-2

-3

.2 .1 0 1 2 3

_c{%}

S&C k:e S_e

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

_J

f

/

-I

5<

-2

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

.31

-2

2 3 4

Ice Shape Tap Locations

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

Next: Appendix B →

Appendix B

Appendix B Run Log Test 09-20-94 Start of Test 09-01-94, End of Velocity Elevator AOA /Remarks Run Configuration (kts) (deg) (deg) 60.0 0.0 18,16,14,12,10,8,6,4,2,0,-2,-4,-6,-8, 01 Baseline -10,-12,-14,-16,-18 " 100.0 0.0 12,8,4,0,-4,-8,-12,-14,-16,-18 -10.0 03 " 60.0 tl 04 " 100.0 " 60.0 - 18,-20,-22 O5 -20.0 12,8,4,0,-4,-8,- 12,- 14,- 16,- 18 If " 100.0 Check Run O8 Check Run -20.0 Baseline 60.0 12,8,4,0,-4,-8,-12,-14,-16,-18 l0 " 100.0 ll " 60.0 -26.6 12,8,4,0,-4,-8,- 12,- 14,- 16 II " 100.0 12,8,4,0,-4 -26.6 14 " 100.0 -8,-12,-14,-16 10.0 " 60.0 12,8,4,0,-4,-8,- 12,- 14,- 16,- 18 " 100.0 " 60.0 14.2 II " 100.0 14.2 12,8A,0,-4,-8,- 10,- 12,- 14, S&C Ice Shape 60.0 -16,-18 _t " 100.0 8,4,0,-4,-8,-10,-12,-14,-16 10.0 " 60.0 8,4,0,-4,-8,- 12,- 14,- 16 " 100.0 " 60.0 0.0 8,4,0,-4,-6,-8,- 10,- 12,- 14 wt ii " 100.0 vv " 60.0 -10.0 Bad run 26 " 100.0 " 100.0 8,4,0,-4,-6,-8,- 10,- 12,- 14 -20.0 " 60.0 8,4,0,-2,-4,-6,-8,- 10,- 12 11 Tt " 100.0 T* " 60.0 -26.6 ft " 100.0 -26.6 LEWICE Ice Shape 60.0 *v 33 " 100.0 t* -20.0 " 60.0 tt " 100.0 Bad run NASA/CR--2000-209921 ]VOL 1 47 AOA /Remarks Run Configuration Velocity Elevator (kts) (deg) (deg) 37 LEWICE Ice Shape 60.0 - 10.0 8,4,0,-2,-4,-6,-8,- 10,- 12 38 " 100.0 " 39 Not Used 40 " 60.0 0.0 8,6,4,2,0,-2,-4,-6,-8,- 10, -12,-14 w 41 " 100.0 " 42 " 60.0 10.0 8,4,0,-4,-8,- 10,- 12,- 14,- 16 *t 43 " 100.0 " 44 " 60.0 14.2 8,4,0,-4,-8,- 10,- 12,- 14,- 16 45 " 100.0 " 46 ......

Repeat of 045 47 " 100.0 0.0 Repeat of 041 48 .... -26.6 Repeat of 033 49 ......

Repeat 2 of 033 50 Baseline ....

Repeat of 013 51 ......

Repeat 2 of 013 52 .... -10.0 Repeat of 004 53 " 60.0 " Repeat of 003 54 " 60.0 0.0 Repeat of 001 55 " 100.0 " Repeat of 002 56 " 100.0 " Repeat of 055 57 " 100.0 14.2 Repeat of 018 58 S&C Ice Shape 60.0 " Repeat of 019 59 " 100.0 " Repeat of 020 60 .... 0.0 Repeat of 024 _ _v In Flow Viz 61 .... -26.6 Repeat of 031, Flow Viz Baseline 60.0 " Flow Viz " 60.0 0.0 Flow Viz 62 Baseline, NP 100.0 " 16,12,8,4,0,-4,-8,- 12,- 14,- 16, - 18,-20, Flow Viz 63 Section 2, NP 60.0 " 8,4,0,-4,-8,- 10,- 12,- 14,- 16,Flow Viz tu 64 .... -20.0 ii 65 .... 14.2 66 .... 0.0 - 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 - 0001,0002,0003,0023,0024,0040,0041.0063 4. 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

Next: Appendix C →

Appendix C

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 0011 Run 0012

Run 0010

Date09/12/94 Date09/12/94

Date09/12/94

AOA Time AOA Time AOA Time

12 1758

12 1553 12 1639

8 1557 8 1643 8 1801

4 1647 4 1805

4 1600

0 1650 0 1808

0 1605

-4 1812

-4 1606 -4 1653

-8 1610 -8 1656 -8 1815

-12 1613 -12 1700 -12 1818

-14 1704 -14 1182

-14 1616

-16 1707 -16 1825

-16 1620

-18 1711

-18 1623

Run 0015 Run 0016

Run 0013/14

Date09/12/13/94 Date09/13/94 Date09/13/94

AOA Time AOA Time

AOA Time

12 1015 12 ll01

12 1836

8 1105

8 1840 8 1019

4 1108

4 1843 4 1022

0 1847 0 1026 0 1111

-4 1850 -4 1029 -4 1115

-8 1032 -8 1118

-8 0929

-12 1035 -12 1122

-12 0932

-14 1038 -14 1125

-14 0936

-16 1042 -16 1128

-16 0939

-18 1045

-18 1745

Run 0018 Run 0019

Run 0017

Date09/13/94 Date09/14/94

Date09/12/13/94

AOA Time AOA Time

AOA Time

12 1305 12 1006

12 1220

8 1010

8 1224 8 1309

4 1013

4 1228 4 1312

0 1017

0 1232 0 1316

-4 1235 -4 1319 -4 1020

-8 1322 -8 1024

-8 1238

-12 1326 -12 1027

-12 1241

-14 1030

-14 1244 -14 1329

-16 1037

-16 1247 -16 1332

-18 1251 -18 1336 -18 1040

NASAJCR--2000-209921/VOL 1 50 Run 0020 Run 0021 Run 0022 Date 09/14/94 Date 09/14/94 Date 09/14/94 AOA Time AOA Time AOA Time 8 1110 8 1216 8 1256 4 1114 4 1300 4 1220 0 1118 0 1223 0 1304 -4 1121 -4 1227 -4 1307 -8 1124 -8 1230 -8 1310 -10 1127 -10 1233 -10 1313 -12 1131 -12 1236 -12 1317 -16 1320 -14 1135 -16 1240 -16 1138 -18 1243 -18 1323 Run 0024 Run 0025 Run 0023 Date 09/14/94 Dme 09/14/94 Date 09/14/94 AOA Time AOA Time AOA Time 8 1400 8 1458 8 1628 4 1406 4 1505 4 1633 0 1635 0 1414 0 1509 -4 1639 -4 1415 -4 1515 -6 1421 -6 1520 -6 1642 -8 1428 -8 1525 -8 1645 -10 1430 -10 1530 -10 1649 -12 1435 -12 1537 -12 1652 -14 1655 -14 1441 -14 1541 Run 0028 Run 0029 Run 0026/27 Date 09/14/94 Date 09/14/94 Date 09/14/94 AOA Time AOA Time AOA Time 8 1708 8 1929 8 2010 4 2013 4 1711 4 1932 0 2016 0 1715 0 1938 -2 2020 -4 1718 -2 1939 -6 1722 -4 1943 -4 2023 -8 1725 -6 1946 -6 2026 -10 1728 -8 1949 -8 2029 -10 2033 -12 1731 -10 1953 -14 1849 -12 1956 -12 2036 NASAJCR---2000-20992 I/VOL 1 51 Run 0032 Run 0030 Run 0031 Date 09/15/94 Date 09/15/94 Date 09/15/94 AOA Time AOA Time AOA Time 8 0955 8 1455 8 0910 4 1459 4 0913 4 1000 0 0918 0 1005 0 1503 -2 0921 -2 1007 -2 1506 -4 1010 -4 1510 -4 0923 -6 1013 -6 1513 -6 0927 -8 1017 -8 1516 -8 0930 -10 0934 -10 1021 -10 1519 -12 0937 -12 1023 -12 1522 Run 0034 Run 0035 Run 0033 Date 09/15/94 Date 09/15/94 Date 09/15/94 AOA Time AOA Time AOA Time 8 1537 8 1658 8 1832 4 1702 4 1836 4 1541 0 1705 0 1839 0 1545 -2 1708 -2 1843 -2 1548 -4 1711 -4 1846 -4 1551 -6 1850 -6 1554 -6 1715 -8 1558 -8 1718 -8 1853 -10 1601 -10 1722 -10 1856 -12 1605 -12 1725 Run 0040 Run 0037 Run 0038 Date 09/16/94 Dme 09/15/94 Date 09/15/94 AOA Time AOA Time AOA Time 8 2012 8 2053 8 0904 4 2015 4 2056 6 0910 0 2100 4 0915 0 2018 2 0920 -2 2021 -2 2103 -4 2025 -4 2107 0 0925 -2 0930 -6 2028 -6 21t0 -8 2114 -4 0934 -8 2031 -10 2117 -6 0940 -10 2034 -12 2037 -12 2120 -8 0945 -10 0950 -12 0955 -14 1000 NASA/CR--2000-209921]VOLI 52

Run 0041 Run 0042 Run 0043

Date09/16/94

Date09/16/94 Date09/15/94

AOA Time AOA Time

AOA Time

8 1610

8 1159 8 1400

4 1614

6 1205 4 1405

4 1210 0 1409 0 1617

2 1215 -4 1412 -4 1620

-8 1415 -8 1624

0 1220

-10 1418 -10 1627

-2 1225

-4 1230 -12 1421 -12 1630

-6 1235 -14 1425 -14 1633

-8 1241 -16 1428 -16 1637

-10 1245

-12 1250

-14 1256

Run 0045

Run 0044

Date09/15/94

Date09/16/94

AOA Time AOA Time

8 1730 8 1828

4 1735 4 1833

0 1835

0 1738

-4 1741 -4 1838

-8 1745 -8 1842

-10 1748 -10 1845

-12 1752 -12 1848

-14 1755 -14 1852

-16 1855

-16 1758

NASA/CR--2000-20992 I/VOL 1 53

Next: Appendix D →

Appendix D

Appendix D Solid 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, AV a (The subscript u indicates uncorrected data) V,, tl: 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 C E.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: 1 c AO_ = 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 with: C,, = Ch,,(1- 2e )+-4-CI,- 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

Next: Appendix E →

Appendix E

Appendix E Data Reduction Output File Formats Uncorrected Coefficients File Name: AA500XX.DAT Run: 00XX Column Value AOA (degrees) Elevator Deflection (degrees) 3 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 AOA (degrees) Elevator Deflection (degrees) 3 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 AOA (degrees) Elevator Deflection (degrees) 3 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 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 AOA (degrees) Elevator Deflection (degrees) Mach Number Reynolds Number (millions) 5 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

Next: Appendix F →

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

NASA/CR--2000-209921/VOL 1 63

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NASA/CR--2000-209921/VOL 1 102 Form Approved

REPORT DOCUMENTATION PAGE

OMB No. 0704-0188 Public reporting burden for 1his collection of information is estimated to average 1 hour per response, includingthe time for reviewinginstructions, searching existing data sources, gathering and maintaining the data needed, and completing and reviewing the collection of informalion. Send comments regarding this burden estimate or any other aspect of this collection of information, including suggestions for reducing this burden, to Washington Headquarters Services, Direclorate for Information Operations and Reports, 1215 Jefferson Davis Highway, Suite 1204, Arlington, VA 22202-4302, and to the Office of Management and Budget, Paperwork Reduction Project (0704-0188), Washington, DC 20503 1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3, REPORT TYPE AND DATES cOVERED 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 WU-548-21-23--00 NAG3-1574 6. AUTHOR(S) Dale Hiitner, Michael McKee, Karine La Nor, and Gerald Gregorek 8. PERFORMING ORGANIZATION 7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES) REPORT NUMBER Ohio State University Research Foundation E-12159-1 1960 Kenny Road Columbus, Ohio 43210-1063 9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES) 10. SPONSORING/MONITORING AGENCY REPORT NUMBER National Aeronautics and Space Administration John H. Glenn Research Center at Lewis Field NASA CR--2000-209921 -VOL1 Cleveland, Ohio 44135-3191 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 12b. DISTRIBUTION CODE Unclassified - Unlimited Subject Categories: 02 and 05 Distribution: Nonstandard This publication is available from the NASA Center for AeroSpace Information, (301) 621-0390.

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 15. NUMBER OF PAGES Aircraft icing; Tailplane icing; Airfoil performance 16. PRICE CODE

AQO

17. SECURITY CLASSIFICATION 18. SECURITY CLASSIFICATION 19. SECURITY CLASSIRCATION 20. LIMITATION OF ABSTRACT OF REPORT OF THIS PAGE OF ABSTRACT Unclassified Unclassified Unclassified NSN 7540-01-280-5500 Standard Form 298 (Rev. 2-89) Prescribed by ANSI Std. Z39- ! 8 298-102

Overview

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7.7 5-Hole Probe Results

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Appendix A View in PDF

Section 2, Clean View in PDF

S&C Ice Shape Taps

TAP # X/C Y/C

1 -.0084 .0030

2 -.0027 .0201

3 -.0014 -.0150

4 -.0018 .0315

5 -.0019 -.0264

6 .0088 .0308

7 .0088 -.0275

8 .0175 .0286

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LEWICE IceShape Taps

TAP # X/C Y/C

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2 -.0207 .0338

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4 -.0044 .0239

5 -.0044 -.0213

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

Section 1 (Baseline) View in PDF

Section 1 (Baseline)

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Appendix A

Section 2, Clean

Section 1 (Baseline)

Appendix B

Appendix C