Large-scale wind-tunnel tests of inverting flaps on a STOL utility aircraft model

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NASA TP

1696 1

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NASA AVRADCOM

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Technical. Paper 1696 Technical Memorandum 80-A-1 '

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Large-Scale - Wind-Tunnel Tests

of Inverting Flaps . on a STOL

Utilitv Aircraft Model.

d P. Morelli Terrell W. Feistel and Joseph J U N E 1980 TECH LIBRARY KAFB. NM

AVRADCOM

NASA

TechnicalMemorandum 80-A-1

Technical Paper 1696

Large-Scale Wind-Tunnel Tests

of Inverting Flaps on a STOL

Utility Aircraft Model

Terrell W. Feistel Ames Research Cerlter Moffett Field, Califorlliu Joseph P. Morelli Aeromechar~ics Laborutory A V R A D C O M Research urd TechnologyLaboratories Ames Research Center Moflett Field, California National Aeronautics , and Space Administration Scientific and Technical Information Office ' " . . . , .

NOTATION b wing span - C mean aerodynamic chord C wing chord drag coefficient, - D CD 4s flap chord cF flap hinge-moment coefficient about flap pivot, ~ HMF cHF @FCF L lift coefficient, - CL 4s M pitching-moment coefficient at 0.25 c , - c m 4 s c D drag flap hinge-moment about flap pivot ffMF horizontal tail incidence relative to wing chord plane, deg i f f L lift M pitching moment free-stream dynamic pressure wing area S flap area per panel SF T thrust T thrust coefficient, - 4s CY angle of attack ofwing chord plane, deg flightpath angle, deg flap deflection, deg/vane deflection, deg inboard outboard iii LARGE-SCALE WIND-TUNNELTESTS OF INVERTING FLAPS ON A STOL UTILITYAIRCRAFT MODEL Terrell W. Feistel Ames Research Center and Joseph P. Morelli Aeromechanics Laboratory AVRADCOM Research and Technology Laboratories SUMMARY A unique inverting flap system has been investigated on a large-scale deflected slipstream model in the Ames 40- by 80-Foot Wind Tunnel. The term “inverting” isused because the flaps pivot about a point near the wing trailing edge and are retracted into the wing contour at a deflec- tion approaching 180”. The subject tests utilized 33% chord double-slotted flaps on a low-aspect ratio wing that was fully immersed in the propeller slipstream. Evaluation of the flap effectiveness is aided by comparisons with the results of tests of other flap systems on the same twin propeller, twin tail boom STOL utility aircraft model.

No extreme or abrupt force or momentincrements were encountered when theflaps were deflectedthrough a wide range, corresponding tothe completeretraction/extensionspectrum.

Hinge-moment informationforthemoreimportant flap deflections was obtainedfrominstru- mented flap supports and exhibitedacceptable behavior. Integral spoilers wereinvestigated for lateral control and were found to produce appreciable rolling moments.

The lift and descent capability of the inverting flaps comparedveryfavorably with that of other flap systems that have been tested on this model, including some with much greater mechan- ical complexity. As expected, the flaps caused large nose-down pitching moment increments at the high lift setting; however, the trimmed characteristics are still competitive with those obtained with the more complicated flap systems.

INTRODUCTION The inverting flap concept was originated by Albert0 AlvarezCalderon (ref. l), who designed the flaps used in the tests reported herein. “Inverting” flaps are large, chord extending flaps that - pivot about a point near the wing trailing edge and are retracted into the wing contour at a deflec- tion approaching 180”. One of the principal differences (as elaborated in ref. 1) between inverting flaps and the more commontypesof chordextending flaps isin theirpotentially simpler (and lighter) method of mounting and installation, presumably requiring only a pivot and an extensible cable. Other significant differences include the potential rapid retract and deployment capabilities and thewide range (0" to 180') of settings possible.

The subject wind-tunnel testswere initiated to shed lighton some of the unanswered questions regarding the practical use of these flaps. In particular, the potential problems associated with trim- ming such large chordextending flaps, possible undesirable aerodynamic force and moment tran- sients during deployment and retraction of the flaps, and the magnitude of the hinge moments encountered, were investigated.

DESCRIPTION OF THE MODEL The model with flaps deflected is shown mounted in thewind tunnel in figure 1. It represents, in approximately full scale, a generalized twin propeller, twin tail boom, deflected slipstream STOL utilityaircraft. A three-view sketch of the model is shown in figure 2(a). Figures -2(b)and2(c) show typical airfoil and flap cross sections, and spoiler and end plate details, respectively. The flap and airfoil coordinates are given in table 1. Close-up photographs of the flap segments set at various deflections are shown in figure 3. Further details of the model may be found in reference2.

TESTS Tests to determine thelongitudinal characteristics of the model were made for flap/vane deflec- tions ( 6 ~ v) of 20/9.5,40/21,60/32, 120/13.5, 180/0, and 60/32 inboard - 90/37 outboard, with a nomina I free-stream dynamic pressure of 192 N/m2 (4 psf) and with nominal thrust coefficients of 0, 1 , 2.4, and 4 at most of the deflections. The data are referenced to wind axes for the forces and to stability axes for the moments. Except as noted in the "DiscussionandAnalysis" section, the data are computed about the 25% wingmean aerodynamic chord position on the wing chord line. The nominal Reynolds number based on the wing mean aerodynamic chord of 2.13 m (7.0 ft) was 2.53X lo6.

The data are not corrected for model support strut tares or for wind-tunnel wall effects. For this support strut system, experience has shown that the tares are negligible for high lift configura- tions. As a basis of comparison for tunnel wall effects, a representative calculation was made using the C method(cf. ref. 3, p. 9), in which the classical correctionsare applied to the circulation Laem portion of the lift only (after subtracting an estimate of the reaction lift attributable directly to thrust).Approximatecorrectionsfor the assumed representative maximum performancedescent condition (CL = 6 at y = -lo", T i = 2) are as follows: AC = +0.07 (tail on)

M J

TABLE 1 .- INVERTING FLAP SYSTEM

(a) M a i n wing airfoil ordinates in fraction of wing chord. (Model scale wing chord length c = 2.13 m (84 in.); basic airfoil section NACA 63A-418 (a = 0.8 mod); leading-edge radius 0.0235 c on slope through leading edge of 0.1860.)

Upper surface Lower surface

- x/c - Y / c

d c

ylc 0.0023 0.0153 0.0131 0.0077

.0045 .O 1 87 .o 105 -.0155

.0090 .0243 .O 160 -.0195 .0209 .0349 .029 1 - .0265 .0453 .OS04 .OS47 - .0360

.070 1 .062 1 .0799 - .0426

.095 1 .07 17 .lo49 - .0478

.1453 .0868 .1547 -.0555 .1957 .0979 .2043 - .0607 .2463 .lo60 .2537 - .0640 .2970 .1115 .3030 - .0656 .3477 .1144 .3523 - .0654

.3984 .1148 .40 16 - .0636

.449 1 .1129 .4509 - .0603 .4998 .lo89 .5002 -.0557 .5504 .lo32 .54 - .OS03 .60 10 .0959 .55 - .0484 .65 15 .0872 .56 - .0460 .70 19 .0772 .58 - .0390 .7522 .0664 .60 - .030 1 .8026 .0546 .65 -.0105 .8525 .04 15 .69 .005 1 .87 .0375 .7 1 .O 128.

.88 .0345 .73 -0 197 .89 .03 15 .75 .0249 .90 .0282 .80 .03 19 .9 1 .0249 .85 .0343 .0341

, .86

I Pivot points: I .87

.0334 .03 19 .0297

Vane: .8770 I -.0101 .90 .0270

I I .91 .0240

TABLE 1 .- CONTINUED.

(b) Flap ordinates in fraction of flap chord. (Model scale flap chord length 0.71 m, 0.33 c (28 in.); leading-edge radius 0.035 cflap on chordline.)

0 0 .O 135 .0290

- .0260

.025

.04 10 - .0345

.050 .0550 - .0430

.075 .0640 -.0495 -10 .0701 -.0535

.15 .0760 - .0600

.20 .0765 - .0660

.25 .0745 -.0715

.30 .0690 - .0755

.35 .0645 - .0780

-40 .0590 - .0770 .45 .0540 - .0735

S O .0490 - .0680

.60 .0390 -.0548 .70 .0290 -.0395 .80 .O 190 -.0235

.90 - .0095

.0090

.9 5 .0045 - .0040

1 .o 0 0

Flap pivot point: TABLE 1 .- CONCLUDED.

(c) Vane ordinates i n fraction of vane chord. (Model scalevanechord length 0 . 2 5 m, 0.12 c (1 0 i n . ) ; leadingedge radius 0.038 cVme on chord line.)

0 0

0 . 0 1 2 5 0.0580 - .0360

.0250 .0790

- .0490

.050

.1120 - .0580

.075 . I 340 - .0600

.IO

.1520 - .0580

.15 .I770 -.0510 .20 .1990 - .0480

. 2 5 .1990 - .0420

.30 .2000 -.0355

.3 5 .1950 - .0300

.40

.I810 - .0240

.45 . I700 -.0190

S O .1540 -.0130

.60 .I240 - .0020

.70 .0940 .0090 .80 .0640 .O 185 .85 .0490

.o 190

.90 .0340 .O 140 . 9 5 .O 1 90 .0040

1 .oo .0040 - .0040

Vane pivot point: It is conceded that such corrections (for anextreme condition) arenot negligible. But they are not of such a magnitude as to invalidate the conclusions drawn by examining the uncorrected data.

The data are presented in uncorrected form in order to be directly comparable withthe previously published dataforthe same and similar models(ref.2). Wall correctionscan be applied to the data using the classical method as follows:

CD = CD + 0.0076 CL

U U and, with the tail on

C ' = CM + 0.0175 CL

U U where the subscript u indicates data uncorrected forwall effects.

Propeller thrust was determined by taking the difference between longitudinal force measure- ments with propellers operating and with propellers removed. Propeller thrust measurements were made at several propeller speeds for each of three free-stream dynamic pressures withtheflaps retracted and with the model at the angle of attack for zero lift. The propellers were contrarotating and propeller rotation was down inboard for all of the tests. Details concerning the propeller design may be found in reference 2.

To obtain flap hinge-moment data, the flap positioning links on the right wing were fitted with in fig- strain gauges. The integral spoiler was installed in the right outboard panel only, as shown ures 2(a) and 2(c) and was deflected 40" from the nesting position for the lateral control effective- ness tests. The horizontal stabilizer was pivoted at the top of the vertical fins and was set at several discreet angles as noted. The integral hinged elevator was set at zero with respect to the stabilizer throughout the tests.

RESULTS An index to the basic data figures is presented in table 2. The basic data are presented without discussion. Figures 4 through 8 present data for a range of flap settings,both tail off and tail on at various stabilizer settings. Figure 9 shows theeffect ofa spanwise variation in flapdeflection (hybridsetting), and figure 10presentsresultsobtained with apartially folded flapposition = 120"/13.5). The planewing characteristics SF/^ = 180"/0) are presented in figure 11.

The effects of spoiler deflection for roll control are shown in figure 12 for two flap settings ( 6 ~ / y = 60/32 and 90/37). The simple spoilers tested are capable of providing more than adequate rolling moments, that are relatively invariant with CL for constant T i .

TABLE 2.- INDEX TO DATA FIGURES " ~~ iHa Remarks &F/ V ~ " . ~ - - 60132 Tail off +5 - 5

1 -15

- Tail off 4012 1 4012 1 +5 - Tail off 2019.5 2019.5 +5 2019.5

- 10

- 90137 Tail off 90137 +5 90137

- 10

9012 1 - 10

60132 I I B - 10 Hybrid setting

90137 011 120113.5 +5 +5 Stored position 18010 60132 - 15 40" spoiler, rt. wing O./B.

(1at.direc. data) 60132 -15 40" spoiler, rt. wing O./B.

(long. data) 40" spoiler, rt. wing O./B.

90137 - 10

(1at.-direc. data) 90137 - 10 40" spoiler, rt. wing O./B.

(long. data) 2019.5 +5 Flap hinge movements (rt.1 4012 1 60132 90137 120113.5

I

aElevator deflection zero with respect to stabilizer, all runs.

The flap hinge-moment data are presented in figure 13. The static hinge moments fall within a reasonable range for all flap deflections, indicatingthat no significant structural problems should be encountered, provided the deployment sequence is programmed to avoid high impact loads. No negative (compressive) loads were detected under any test condition, indicating that simple exten- sible cables should suffice to position the flaps.

DISCUSSION AND ANALYSIS Potential Performance Capabilities Figure 14 shows a representative set of polars for a hypotheticalSTOL airplane, similar to the wind-tunnel model, with inverting flaps set at 60/32. Thewind-tunnel data have been adjusted for a trimming tail load with the c.g. set a t 0.35 c on the thrust line.

Several sets of constant-value linesfor awing loading of 2394 N/m2 (50 psf) are superimposed on the basic polars in this figure. Radiating from the origin are lines of constant descent angle, 7.

The curved dashed linefrom the .oiiginrepresentsa sink rate of 6.10m/s(20fps), which isa nominal landing gear limitforthis type of aircraft. For reference,horizontal lines of constant approach speed are shown. The hatched lines are approximate boundaries of constant “no-flare” landing distance’ over a 15.2-m (50-ft) obstacle.

A shaded circle is superimposed to indicate a potentially feasible landing approach condition for a 152-m (500-ft) landing. This cokesponds to an approach speed of 50 knots (CL = 5.9) and a descent angle of approximately 10” (Co = 1.03).Thethrustcoefficient would approximate 2.0 (approximately 800 total S H P (59,656. W)), and the angle ofattack would approximate15” (nose-up attitude of 5”) for the hypothetical STOL airplane.Theapproachrateof sink in this condition would be approximately 4.5 m/s (1 5 fps).

. Comparison With OtherFlapSystems Figure 15 compares the relative descent performance for three different high lift systems: a simple double-slotted flap; a rotating cylinder flap; and the inverting flaps. All three fold into the same basic wing contour for cruise and have been tested in the Ames 40- by 80-Foot Wind Tunnel using this sameSTOL utilityaircraftmodelwith the same basic wing planform and propulsion system. The rotating cylinder flap system is the most mechanically complicated,’being a form of mechanical boundary-layer control, and is described in reference 2.

As a basis for comparison (assuming a 2394 N/m2 (50 psf) wing loading throughout), a repre- sentative maximum usable descent condition is shown for each flap system. The descent angle of * The hypothetical “no-flare” landing distances are based on a straight approach path over a 15.2-m (50-ft) obstacle to touch- down, followed by a constant 0.5 g linear deceleration to stop. They are meant only as a guide to potential maximum performance STOL landing capability, which can be easily superimposed on the polar plot. The “no-flare” landing distances, as calculated, corre- spond very closely to flight-test data for actual “half-flare” landings i n current generation STOL aircraft, as shown in reference 4; they serve, therefore, as a valid basis for comparison.

attack is arbitrarily chosenas 10" below the angle of attack formaximum lift. Shown superimposed are the constant, no-flare landing distance lines introducedin the previous figure. Therotating cylinderflap and the invertingflapare seen to be quite evenly matched ata descent angleof approximately 13" to 14",corresponding to atheoretical "no-flare" landing distance of about 122 m (400 ft) over a 15.2-m (50-ft)obstacle.Thedouble-slottedflap system (withmuch less chordextension) has about one-half this maximum descentcapability,withadescent angleof 7" to 8", corresponding to a distance of about183 m (600 ft).

Summary of Longitudinal Data for a Range of Flap Settings Figure 16(a) summarizes, for T i = 1 .O, longitudinal data (at iH = +5") covering the complete range of flap settings tested..It can be seen that no extremely abrupt force or moment changes are encountered in going from the 180" folded position to one of the lifting positions (20" to 90") and back again.2 Figure16(b) shows, forreference,apitchingmomentplot of the same data recomputed with the moment reference located at 0.25 Flongitudinally but, more appropriately, on the thrustline instead of on the wing chord. It will be noticed that some of the pitching moment curves have their positions shifted with respect to each other, but no large effects are involved.

Longitudinal Stability and Trim Considerations In order to evaluate more easily the longitudinal stability and trim considerations involved in using such highly effective flaps, the data have been recomputed with respect to a c.g. on the thrust lineat the 0.35 c point. This is amorerepresentativelocationforadeflected slipstream STOL aircraft.

Figure 17(a) is a summary plot with the moment reference at 0.35 c on the thrust axis. It shows the pitching moment curves for the complete range of flap settings with T i = 1 and i H = +5".

(Data were available for all flap settings at i H = + S o . More nose-up trim is, of course, available at the more negative stabilizer settings.) It can be seen that, except for the 6 ~ = 120" transition position, all flap settings exhibit positive or neutral stability throughout the usable angle-of-attack range for these conditions.

Tofacilitate analyses withreference to this c.g. location, figures 17(b)through17(e) show representativepitching-momentcharacteristicsfor flap deflectionsof 60/32 ( i H = -So), 90/37 ( i H = -lo"), 60/32 inboard - 90/37 outboard (iH= -lo"), and 20/9.5 ( i ~ = -lo"), covering a wide rangeof thrustcoefficientswith the reference at 0.35 c on the thrust axis. Figure 17(0 shows representative tail-off pitching-momentdataaboutthisreference at T,' = 2.4 forthreeflap settings: 60/32,90/37, and 20/9.5.

It is obvious that the horizontal tail used on the wind-tunnel model was not adequate to trim these flaps at their most effective settings. In fact, itwas far from optimum for ahigh-performance STOL aircraft; it was relatively small, poorlyendplated, and had a simple nonslottedcontrol surface (whichwas not deflected for this test).

It is anticipated that the SF = 120" position would not be used for steadystate flight, but that the flap would be moved continuously through such transition positions to arrive at the more effective deflections.

A preliminary study indicated that the standard horizontal tail of the OV-1OA aircraft would be adequate for trimming the pitching moments produced by the 6~ = 60/32 inverting flap con- figurationreportedherein up to a T,‘ of 2.4, witha maneuvering reserve capabilityof about 0.6 rad/sec2. A slightly larger horizontal tail would be required to provide adequatetrimand maneuver capability up to T i . = 4.

CONCLUDING REMARKS Tests havebeen performed on a full-scale, twin-boom deflected slipstream STOL utility air- craft configuration equipped with “inverting” flaps in the Ames 40- by 80-Foot Wind Tunnel. The results show that the lift and descent capability provided by these flaps compares favorably with that of other flap systems that have been tested on the same model, including some with much greater mechanical complexity.

In spite of a large tail-off nosedown pitching moment produced by the inverting flaps in the highest lift configurations, the trimmed characteristics are competitive with those obtained from the more complicated flap systems. The flap hinge moments exhibited acceptable behavior throughout the range of deflections tested.

It is believed that these flaps may have promising potential application to the design of rela- tively simple STOL utility aircraft with improved performance capabilities. In addition, they may merit consideration as retrofits to existing aircraft withless effective flap systems.

Ames Research Center National Aeronautics and Space Administration Moffett Field, Calif. 94035, February 1980 REFERENCES 1. Alvarez-Calderon, Alberto: Design andTests of Inverting Flapsand Wing Span Flaps. S.A.E.

Paper No. 680646 presented at the Aeronauticsand Space Engineering andManufacturing Meeting, Los Angeles, Calif., Oct. 1968.

2. Weiberg, J . A.; and Gamse, Berl: Large-Scale Wind Tunnel Tests of an Airplane Model with Two Propellers and Rotating Cylinder Flaps. NASA T N D-4489, 1978.

3. Koenig, D. G.;and Corsiglia, V. R.: Aerodynamic Characteristics of a Large-ScaleModel with an Unswept Wing and Augmented Jet Flap. NASA TN D-4610, 1968.

4 . Feistel, T. W.; and Innis, R. C.: Results of a Brief Flight Investigation of a COIN-TypeSTOL Aircraft, NASA TN D4141, 1967.

(a) Threequarter rearview, 6 ~ / v = 60/32.

Figure 1 .- Photos of model mounted in tunnel.

(b) Threequarter front view, 6 ~ / v = 60/32.

Figure 1 .- Concluded.

Wing Area 16.25 m2 (175 f t 2 ) Span 7.62 m (25.0 ft) Chord 2.13 m (7.0 f t ) Aspect r a t i o 3.57 Section 63A-418 (mod.)

Horizontal tail Area 5.60 m2 (60.3 ft2) Span 4.43 m ( 1 4 . 5 4f t ) Chord 1.26 m (4.15 f t ) Aspectratio 3 30 Section 63-216 (inv.) , Tal I length 5.65 m (18.56 f t ) right wing only 0.86 m Reference Moment center on wing ( 2 . 8 3f t ) Wing chordlineat 0.25 chord chnrd

2.83 m (9.3 f t ) ' \

3 blades Thrust *I AF-IZl/blade axis ( 2 5 . 0 7.62 ml- f t ) (a) Three view.

Figure 2.- Geometry of model.

Vane pivot Vanepivot (b) Typical airfoil and flap cross section and arrangement, 6 ~ / v = 60/32 and 6 ~ / v = 180/0.

Figure 2 .- Con tinued.

1 . 3 6 ~ I. 17c 'Path o f f l a p T.E.

( For BF = 60") ( R a d i u s = 0 . 4 l c )

1 - . 0 . 6 0 c -

I

O.'36c ( c ) Detail of spoiler and end plate geometry, 6 ~ 1 ~ = 60/32 shown.

Figure 2.- Concluded.

6 F/v= 90137 SFIv = 2019.5 Figure 3 .- Photos of several flap deflections.

C L I I O L 1 -3 -2 - I 0 I 2 3 -10 0 IO 20 30 0 -A -.8 -1.2 -1.6 Crn CD 0 , deg (a) Horizontal tail removed.

Figure 4.- Basic longitudinal data, 6 ~ , v = 60/32.

-3 - - I

0 I 2 -.4 -1.2 -1.6 CD Figure 4.- Continued.

I I I 0 I I -10 0 I O 20 30 -3 - 2 - I 0 I 2 3 CD (c) i H = - 5 .

Figure 4.- Continued.

h)

T T

I I I -3 - 2 "I 0 I 2 0 3 -10 I O 20 Q , deg CD (d) i H = -1 5 (stabilizer-elevator juncture taped, lower surface).

Figure 4.- Concluded.

r-

-

-

-

CL

-

- -

-

-

- 4 -3 - 2 - I 0 I 2 -10 0 IO 20 30 0 -.4 -.8 -1.2 -1.6 - 2 . 0 (a) Horizontal tail removed.

Figure 5.- Basic longitudinal data, 6 ~ / v = 40/21.

!

E - 2 - I 0 I 2 -10 0 IO 20 30 4 0 -.4 -.8 CD ' a , deg C m (b) i ~ = +5.

Figure 5 . - Concluded.

r

- -r

r

U I -.4 -.e -1.2 0 I O 20 30 40 0 a , deg Cm (a) Horizontal tail removed.

Figure 6.- Basic longitudinal data, 6 ~ / v = 20/9.5.

N w

.I, .,,, 11111.1 . . . . ... ..-

1 I I I -.4 -.8 -1.2 (b) i H = +5.

Figure 6.- Continued.

CL I - I -4 -3 -2 - I 0 I -10 0 I O 20 30 40 1.2 .8 .4 0 -.4 CD a, deg Cm (c) i H = -10.

Figure 6.- Concluded.

-

(Tail o f f ) TC’ T

E C L i I C - I 0 I 2 -10 0 I O 20 30 40 3 - a , deg CD (a) Horizon tal tail removed, Figure 7.- Basic longitudinal data, 6 ~ / v = 90/37.

-2 - I 0 I 2 - -3 (b) i H = +5.

Figure 7.- Continued.

9- I 8 - 7 - 6- 5 - CL 4 - 3 - 2 - I - 0- -3 -2 - I 0 I 2 -10 0 IO 20 30 Q, de!J CD (c) i H = -10.

Figure 7.- Concluded.

-10 0 IO 20 30 40 .8 .4 0 "4 " 8 Q, deg Cm Figure 8.- Basic Longitudinal data, 6 ~ / v = 90/2 1 ; i H = -1 0.

w

r~

T

T

TCI Swv=60/32 1.8. 0 0 90137 O.B. 0 I 0 2.4

I A 4 I I

-I 0 I 2 3 -10 0 IO 20 30 40 -4 0 -A 7 8 -1.2 CD a , deg Cm Figure 9.- Basic longitudinal data, hF/v= 60/32 inboard, 90/37 outboard; iH= -10.

0 - -3 -2 - I 0 I .4 0 -.4 -.e CD

' Figure 10.- Basic longitudinal data, 6 ~ / v = 120/13.5; iH= "5. "

CD Q, deg Crn Figure 1 1 .- Basic longitudinal data, 6 ~ / v = 180/0 (flap stored); iH = +5.

i H = -15 [Toped) T i 0 0 8 ~ / ~ = 6 0 / 3 2 0 1 GSPLRR.H., 0/0 = 40° A 4 - . 3 -.2 -.I 0 .I . 2 - . 2 -.I 0 .I .2 .3 0 .I .2 .3 .4 CY Cn = I (a) 6 ~ / v = 6 0 / 3 2 ; i ~ = - 1 5 , taped.

Figure 12.- Lateral control effectiveness; spoiler deflected 40°, right hand outboard section.

w w w P i H = -15 (Taped) Tc' 0 0

T

0 1 0 2 . 4 ' A 4 (a) Concluded.

Figure 12.- Continued.

t

Td 0 0 0 1 0 2 . 4 A 4 L -.3 - . 2 - . I 0 . I .2 -.2 - . I 0 . I . 2 .3 -*I 0 .I - 2 -3 CY Cn =t .

Figure 12.- Continued.

3 0 I O 20 30 40 = I deg (b) Concluded.

Figure 12.- Concluded.

7 - Inboard flop T i 0 0 6 - 1.0 2.4 5 -

d

4 - CL 3 - I 2 3 4 5 CHF (a) 6 ~ 1 ~ = 2019.5.

Figure 13.- Flap hinge moment data, right wing.

Inboard Ouiboard flap flap T ; 0 0 0 0 1 . 1 CL I I I I I I 2 3 4 5 C b (b) 6 ~ / T/ = 40/2 1.

Figure .13 .- Continued.

IO Inboard Outboard flap flap T i 0 0 0 0 1 . 1 0 2.3 A n 3.9

;P

CL d; I I I 1 4 5 6 7 ( c ) 6 ~ / v = 60/32.

Figure 13.- Continued.

Inboard Outboard flap flap T i 0 0 0 rn 0 1 . 1

+ 0 2.4

A A 4 I 2 3 4 5 6 CH, (d) 6 ~ / v = 90137.

Figure 13.- Continued.

Inboard flop T , ' 0 0 1 . 1 2.4

I I I I I 1

0 I 2 3 4 5 CHF (e) 6 ~ 1 ~ = 120;13.5.

Figure 13.- Concluded.

V, knots 6 0 TO

-

- .e -1.6 0 .e 1.6 2.4 CD Figure 14.- Trimmed polars with SF/^ = 60/32 for STOL utility aircraft - showing representative landing approach parameters.

152 m (500 f t ) 8~ Z 60 366 m ( 1 2 0 0 f t ) Approximate maximum descent capability: Inverting flap

m Rotating cylinder flap

A Double slotted flap .4 .8 1.2 1.6 2.0 2.4 2.8 3.2 CD Figure 15.- Landing approach condition, descentcapability comparison - three different flap systems.

T

CL I 0 60/32 A 90/37 h 12OA3.5 n 18010 I I I I I I (a) Lift and drag data, cg at 0.25 C, on chord line.

Figure 16.- Summary of TL = 1 longitudinal data for a range' of flap settings.

C L I 0 60/32 A 90137 h 120/13.5 C I I - I -A -.8 - 1.2 Cm0.25i. C. L.

(b) Pitching-moment data, cg at 0.25 E , on chord line.

Figure 16.- Continued.

I I I I - .4 -.8 -1.2 Cm0.25e,T.A.

(c) Pitching-moment data, cg at 0.25 E , on thrust line.

Figure 16.- Concluded.

. . .

. .

T

8F/V 2019.5 40/21 60/32 90/37 120113.5 180/0 I 1 1 . 8 .4 0 - . 4 - . e - 1 . 2 C m 0 . 3 5 E , T. A.

(a) Superposition of T i = 1 data for arange of flap settings, i H = +5" Figure 17.- Pitching-moment data replotted for cg at 0.35 c on the thrust line.

0 2 A 4

I

(b) SF/^/ = 60132, ‘H = -5

Figure 17.- Continued.

I O CL Tc’ 0 0 I n~ 0 2 . 4 A 4 I I I I I C

-A -. 8 -1.2

! .8 .4 ( c ) aFlV = 90/32, i H = -10.

Figure 17.- Continued.

IC C L Tc' 0 0 I 0 0 1 0 2 . 4 A 4

I

OL 1 . 2 .8 .4 0 - .4 -. 8 - 1.2 T.A.

( d ) 6 ~ / v = 60/32 I./B., 90/32 O./B.; iH= -10.

Figure 17.- Continued.

CL I C I I 1 I I I - I -.8 -1.2 I .8 .4 0 -.4 Cm0.35€, T. h .

(e) S F / V = 20/9.5, iH= -10.

Figure 17.- Continued.

I ? ’

b

F/V 0 20/9.5 0 60/32 0 9 W 3 7 I I -.4 ..e -1.2 -1.6 C m 0 . 3 5 F . T. A .

- (f) Superposition of tail-off data for a range of flap settings, T i = 2.4.

Figure 17.- Concluded.

I I1 I I 1 Ill I 1 I 1. Report No. NASA TP-1696 3. Recipient'sCatalog No. 2. Government Accession No.

AVRADCOM TM 80-A-1 4. Title and Subtitle 5. Report Date June 1980 LARGE-SCALE WIND-TUNNELTESTS OF INVERTING 6. PerformingOrganization Code FLAPS ON A STOL UTILITY AIRCRAFT MODEL 7. Authods) 8. PerformingOrganizationReport No.

Terrell W. Feistel and Joseph P. Morelli A-706 1 10.WorkUnitNo.

9. Performing Organization Name and Address 505-10-12 NASA Ames Research Center and 1 1 . ContractorGrantNo.

Aeromechanics Laboratory AVRADCOM Research and Technology Laboratories Moffett Field, CA 94035 13. Typeof Report andPeriod Covered , 12. Sponsoring Agency Name and Address Technical Paper National Aeronautics and Space Administration 14. Sponsoring Agency Code Washington, DC 20546 and U.S. Army Aviation Research and Development Command St. Louis, MO 63166 1 5 . Supplementary Notes Terrell W. Feistel: Ames Research Center Joseph P. Morelli: Aeromechanics Laboratory, U.S. Army R&T Laboratories (AVRADCOM) 16. Abstract A unique inverting flap system has been investigated on a large-scale deflected slipstream model in the Ames 40- by 80-Foot Wind Tunnel. The subject tests utilized 33% chord double- slotted flaps on a low-aspect ratio wing that was fully immersed in the propeller slipstream. Eval- uation of the flap effectiveness is aided by comparisons with the results oftests of other flap systems on the same twin propeller, twin tail boomSTOL utility aircraft mode.

No extreme or abrupt force or moment increments were encountered when the flaps were deflected through a wide range, corresponding to the complete retraction/extension spectrum. The lift and descent capability of the inverting flaps compared very favorably with that of the other flap systems that have been tested on this model, including some with much greater mechanical complexity. As expected,the flaps causedlargenose-down pitching-momentincrements at-the high lift settings; however, the trimmed characteristics are still competitive with those obtained from the more complicated flap systems.

It is believed that these flaps may have promising potential application to the design of rela- tively simple STOL utility aircraft with improved performance capabilities. In addition, they may merit consideration as retrofits to existing aircraft with less effective flap systems.

1 7 . Key Words (Suggested by Author(s)) 18. DistributionStatement High lift flaps Unclassified - Unlimited STOL aircraft Aerodynamics STAR Category - 05 Large-scale testing 19. Security Classif. (of this report1 22. Price' 21. No. of Pages 20. Security Classif. (of this p a g e 1 Unclassified Unclassified $5.25 'For sale by the National Technic4 InformationService. Springfield. Virginia 221 61 NASA-Langley, 1980 Postage and Fees Paid National Aeronautics and T H I R D C L A S S BULK R A T E National Aeronautics and .~ Space Administration Space Administration N A S A 4 5 1

Washington, D.C. I [*)

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