Skip to main content

Full text of "Experimental study of an independently deflected wingtip mounted on a semispan wing"

See other formats


'/v 



NASA Technical Memorandum 102842 



y3-7 (yy 



Experimental Study of an Inde 
pendently Deflected Wingtip 
Mounted on a Semispan Wing 



D. M. Martin and L. A. Young 



(NASA-TM-102842) EXPERIMENTAL STUDY OF AN N92-10983 

INDEPENDENTLY DEFLECTED WINGTIP MOUNTED ON A 
SEMISPAN WING (NASA) 58 p CSCL OlA 

Unci as 
G3/02 0048769 



September 1991 



rWNSA 

National Aeronautics and 
Space Administration 



NASA Technical Memorandum 102842 



Experimental Study of an Inde- 
pendently Deflected WIngtip 
Mounted on a Semispan Wing 



D. M. Martin, University of Kansas Center for Research, Inc., Lawrence, Kansas 
L. A. Young, Ames Research Center, Moffett Field, California 



September 1991 



rVIASA 

National Aeronautics and 
Space Administration 

Ames Research Center 

Moffett Field. California 94035-1000 



TABLE OF CONTENTS 

Page 

NOMENCLATURE '^ 

SUMMARY ^ 

INTRODUCTION ^ 

TEST OBJECTIVES • 2 

EXPERIMENTAL APPARATUS AND PROCEDURES 2 

Model Description and Assembly 2 

Test Conditions and Procedures ^ 

Corrections ■ 

RESULTS.... ^ 

Reynolds Number Effects ^ 

Tip Incidence Angle (AG) Effects ^ 

RC10/08Tip - ^ 

RC10/05Tip ' 

Flow Visualization ° 

Tip Aerodynamic Parameters ^ 

Tip Aerodynamic Center ^ 

Lift and Pitching Moment Data ^^ 

USE OF STEADY FLOW EXPERIMENTAL DATA FOR CALCULATION OF FREE-TIP 

MOTION 10 

CONCLUDING REMARKS 13 

APPENDIX A 1^ 

APPENDIX B 21 

APPENDIX C 25 

REFERENCES 33 

nOURES 34 



. ; iii PRECEDfl^JG PAGE BLANK NOT FJLMED 

PAG£_4JL_INTENT10NALLY BUNK '^' "^'"^^^ 



NOMENCLATURE 

wind tunnel breadth 

form drag component of Cq 

induced drag component (resulting from lift) of C^ 

interference component of Cq 

skin-friction component of Cp 

tip drag coefficient, Dj/qSj Cj 

Cj) for aj = 0.0° 

tip lift coefficient, Lj/qS-i-Cj 

variation of tip lift coefficient with angle of attack, 3Cl /daj , rad-^ or deg-' 

Cl for aj = 0.0° 

tip pitching moment coefficient about the tip aerodynamic center 

change in tip pitching moment coefficient with tip pitch rate 

response of pitching moment coefficient to unit step change in q 

tip pitching moment coefficient about quarter-chord point of inboard wing, 

Mj/qSjCj 

C ^ variation of tip pitching moment coefficient with angle of attack, dC ^^ jda-^ , 

rad~^ or deg~l 

C [. . .] response of pitching moment coefficient to unit step change in a 

C tip pitch damping coefficient, 3C^ l^ip-i ^T /^°° ) 



B 


^form 


Cd, 


Cd„ 


Co,, 


^Dj 




Cl, 


Cl„ 




niac 


C", 


CmqL---. 


^m-p 



a 



C^ C^^ for a-r = 



\l 



,o 



'uiQ m-] 



C C^ for C, =0 



V 



PAGE_4V iNiLisliONALLV BLANK pRECtDifiG PAGE BLAiMK NOT FILMQJ 



C„ C^ at x = 0.0 
•"0.0 n^x 

Cy tip inboard chord length, ft 

Dj tip drag force in wind axis system, lb 

H wind tunnel height 

hj tip vertical velocity component caused by plunging motion, ft/sec 

la tip mass moment of inertia about pitch axis, ft-lb sec^ 

Ks tip torsional spring constant, ft-lb 

Lj tip lift force in wind axis system 

Mx tip pitching moment, measured about inboard wing quarter-chord point, ft-lb 

N measured normal force, lb 

q dynamic pressure, psf 

q J pitch rate, rad/sec 

Re Reynolds number based on tip root chord, pV«,cj/|Li 

RM measured rolling moment, ft-lb 

S measured side force, lb 

Sx tip reference area, ft^ 

t2 streamline curvature effect on angle of attack 

Voo tunnel free-stream velocity, ft/sec 

X chordwise direction in wing-fixed coordinate system 

Xac chordwise position of the tip aerodynamic center 

y spanwise wing station, ft 

aj tip angle of attack 

aw wing angle of attack 



VI 



(ACL)gp lift correction for streamline curvature 

(Aa) J angle-of-attack correction for downwash 

(Aa) angle-of-attack correction for streamline curvature 

Ae tip incidence angle relative to inboard wing, aj - a w ^^^ o^ deg 

e velocity increment 

£5(3 velocity correction for solid-body blockage 

E^^ velocity correction for wake blockage 

e mechanical spring pretwist angle, rad or deg 

e J tip incidence angle relative to horizontal reference plane 

[I coefficient of viscosity, slug/ft-sec 

p sea-level air density, slug/ft^ 



Vll 



SUMMARY 

The results of a subsonic wind tunnel test of a semispan wing with an independently deflected tip 
surface are presented and analyzed. The tip surface was deflected about the quarter-chord of the rect- 
angular wing and accounted for 17% of the wing semispan. The test was conducted to measure the 
loads on the tip surface and to investigate the nature of aerodynamic interference effects between the 
wing and the deflected tip. Results are presented for two swept tip surfaces of similar planform but 
different airfoil distributions. The report contains plots of tip lift, drag, and pitching moment for vari- 
ous Reynolds numbers and tip deflection angles with respect to the inboard wing. Oil flow visual- 
ization photographs for a typical Reynolds number are also included. Important aerodynamic param- 
eters such as lift and pitching moment slopes and tip aerodynamic center location are tabulated. A 
discussion is presented of the relationship between tip experimental data acquired in a steady flow 
and the prediction of unsteady tip motion at fixed-wing angles of attack. 



INTRODUCTION 

The Free-Tip Rotor concept was developed to reduce both the tip oscillatory loading of heli- 
copter blades and the rotor power requirement (ref. 1). This configuration consists of a helicopter 
blade with the last 10% of the length allowed to pitch freely as a rigid body about the quarter-chord 
point of the inboard blade. The free-pitching tip is designed to respond to local changes in inflow 
and thus reduce the amplitude of azimuthally varying tip loads. Because of its simplicity, a nonrotat- 
ing semispan wing with an independently deflected tip is ideal for studying the aerodynamic interac- 
tions between the wing and the tip. The deflected tip can be described as a completely separate tip 
surface attached to the inboard wing with minimal spanwise gap, at an angle of incidence that can be 
varied independently. In this simplified model, the tip is constrained and the angle of incidence is 
fixed while the loads are measured in a steady flow. 

Since a detailed knowledge of tip aerodynamics in subsonic flow is required for design purposes, 
a series of tests has been carried out to measure aerodynamic loads on both wing and tip surfaces 
(refs. 2-4). During these tests, the tip planforms were first-generation concepts that used V23010 air- 
foil sections. A new family of tip planforms has been developed in the last few years (ref 5); these 
incorporate advanced-technology airfoils and geometries that are more compatible with desired rotor 
aerodynamic performance. This report contains the results of a wind tunnel test of a semispan wing 
fitted with these advanced tip planforms. 

The present study is part of the continuing research program associated with the Free-Tip Rotor 
concept. The data presented in this report will be used to define the rotor blade design of an 
Advanced Free-Tip Rotor (AFTR) small-scale wind tunnel model. The report contains plots of tip 
lift, drag, and pitching moment coefficients for different values of tip incidence angle with respect to 
the inboard wing. Oil flow visualization results and some important design parameters, such as aero- 
dynamic center location, are briefly discussed, and calculations of zero-angle-of-attack lift and pitch- 
ing moment coefficients are reported. Possible uses of the measured data as applied to oscillating tips 
in unsteady flow are also proposed. 



TEST OBJECTIVES 



The objectives of this test were as follows: 

1 . To measure the lift, pitching moment, and drag of the new tip planforms in subsonic flow, 
and to investigate the effect of the tip incidence angle, measured with respect to the wing (Ae) 

2. To compute the lift and pitching moment slopes, and the aerodynamic center location, for 
each planform 

3. To study flow patterns on upper and lower tip surfaces and thus qualitatively identify the 
induced aerodynamic effects of the inboard wing on the tip 



EXPERIMENTAL APPARATUS AND PROCEDURES 



This test was conducted in the Ames 7- by 10-Foot Subsonic Wind Tunnel. Tip aerodynamic 
forces were measured with a 3/4-in. -diameter, six-component, internal strain-gauge balance. The bal- 
ance was mounted at the quarter-chord point of the inboard wing. Wing aerodynamic loads were not 
measured. 



Model Description and Assembly 

Two new tip configurations were tested; these are subsequently referred to as the RClO/08 tip 
and the RClO/05 tip. The two tips had the same planform shape and dimensions (fig. 1), but different 
airfoil section distributions. Three section profiles were used; they were based on new airfoil designs 
optimized for high lift and low drag in transonic flow. The airfoil surface coordinates are listed in 
tables Al, A2, and A3 in appendix A. Both tips consisted of three separate regions: an inboard area 
swept 10° with 10%-thick airfoils, a transition region, and an outboard portion with thinner airfoil 
sections. The transition and outboard regions were highly swept. The airfoil sections of the RClO/08 
tip decreased in thickness from the 10%-thick RCIO to the 8%-thick RC08 over 18.8% of the tip 
span (fig. 1). On the RClO/05 tip, the thickness decreased from 10% to 5% over 9.4% of the span 
length. The span of both tips was 0.776 ft, and their planform areas were 0.40 ft^. 

The semispan wing consisted of two parts. The first was a solid steel, rectangular inboard portion 
with V23010 airfoil section profiles and a nonlinear twis t distribution. The V230 10 airfoil-surface 
coordinates are listed in table A4, and the wing twist distribution is plotted in figure Al of appen- 
dix A. The second part was a nonmetric balsa-wood balance housing that matched the inboard sec- 
tion profiles, thus extending the wing inboard span. The wing semispan and area, without tip sur- 
faces, were 3.79 ft and 2.60 ft^, respectively. The wing had a semispan aspect ratio of 5.52 and a 
chord length of 0.686 ft, and was mounted vertically in the test section. A sketch of the wing and tip 
showing the internal balance location and wing-fixed coordinate system is shown in figure 2. 



The tips were attached to the wing as follows: Before the tip was mounted onto the steel wing, 
the balance was inserted into a cylindrical sleeve with a shank attached to one end. The balance/ 
sleeve assembly was then partially inserted into the wing at its quarter-chord point. The tip was then 
mounted over the sleeve, and set screws were placed into the shank to set the tip incidence angle 
with respect to the wing. Several set-screw holes were used to allow various tip deflection angles. 

The moment reference center and the pitch axis, about which tip incidence angles with respect to 
the wing were set, lie at the quarter-chord of the wing section at the interface between the balsa- 
wood insert and the tip surface (fig. 2). The wing angle of attack was measured with respect to the x 
axis at this location. Figure 3 shows the complete assembly in the tunnel test section. The span of the 
entire configuration was 4.57 ft; its semispan aspect ratio was 6.96. The tip planform span accounted 
for 17.2% of the total configuration semispan. 



Test Conditions and Procedures 

The test conditions and ranges of wing and tip angles of attack are listed in table 1. The dynamic 
pressure was varied in increments of 5 psf The maximum Mach number achieved was 0.13. As 
shown in table 1, three values of tip incidence angle with respect to the wing were used for the 
RClO/05 tip and four values for the RClO/08 tip. These were determined from the most feasible set- 
screw positions, given mechanical constraints and design considerations. The data were acquired by 
first setting the tip incidence angle relative to the inboard wing and then varying the wing angle of 
attack. Therefore, since Ae was fixed for a given run, the tip geometric angle of attack (ax) varied 
by the same amount as the wing geometric angle of attack (a^)- 

Table 1. Test conditions and ranges of parameters 



Parameter Value 



Dynamic pressure 5 psf < q < 25 psf 

Reynolds number 2.76 x 10^ < Re < 6. 18 x 10^ 

Wing angle of attack 0.0° < aw ^ 16.7° 

Tip angle of attack -5.26° < aj < 1 8.2° 
Tip incidence angle 

RClO/08 AG = 1.5°, -3.5°, -9.0°, -13.0° 

RClO/05 A9 = -2.5°, -8.0°, -10.0° 

The V23010 airfoil coordinates listed in table A4 of appendix A were defined in terms of a refer- 
ence line bisecting the upper and lower surfaces of the trailing edge (fig. 4). The wing angles of 
attack specified throughout this report were measured from this reference line, since it coincides with 
the wing-fixed x axis. Figure 4 shows a view looking inboard at the wing with the tip removed. The 
orientation of the balance axes is also shown. Since the balance axes are rotated with the wing angle 
of attack, the tip lift, drag, and pitching moment were computed in the wind axis system as follows: 



Ly =NcosaY^ +Ssina^ 

Dj =NsinaYv +Scosa^ (1) 



Mj =-RM 



Corrections 



The data presented here have been corrected for tunnel wall and blockage effects. These correc- 
tions are outlined in detail in appendix B. Since the data were acquired at relatively low dynamic 
pressure and load levels, no corrections were applied to account for balance deformations. 



RESULTS 

The data are presented as plots of Clj vs ttj (lift curves), Cq^ vs Ci^^ (drag polars), and C^^ 
vs ax and C^ vs Cl^ (pitching moment curves). These plots are drawn for several Reynolds 
numbers at coi^stant tip incidence angles (AG) and for various tip incidence angles at constant 
Reynolds numbers. Some poor-quality data were omitted from the plots, so the number of curves or 
data points may not correspond from one plot to another for equivalent test conditions or runs. 

For lift curves and drag polars, the data were fitted by first- and second-order least squares, 
respectively. For tip lift data the curves were fitted only to points corresponding to the linear range of 
the lift coefficient: beyond the stall point, only the data points are shown. The accuracy of force and 
moment coefficients was estimated from known balance gauge accuracies and from tolerances asso- 
ciated with other measured parameters, such as dynamic pressure, reference lengths and areas, etc. 
The estimated combined accuracies were calculated as percentages of the reduced data values as fol- 
lows: for Clt, , ±1.84%; for Cj^^ , ±1.84%; and for C^^,±l.ll%. 

Finally, although two tip planforms were tested, no attempt is made here to compare and rate the 
aerodynamic performance of the tips. The test was not carried out with this objective in mind. One of 
the main purposes of this experiment was to determine the effect of the AG parameter for a given 
tip. This approach yields important information on the mutually induced aerodynamic effects 
between the wing and the tip. 

Reynolds Number Effects 

An analysis of the data plotted at constant AG for different Reynolds numbers illustrates some 
important aerodynamic characteristics of the unique type of configuration discussed in this report. 
Figure 5 shows, for the RClO/08 tip, the variation of tip lift coefficient with angle of attack 
(fig. 5(a)) and the related drag polars (fig. 5(b)), for AG = -3.5°. As expected, the lift curve slopes 
(9Ci / 3aT ) do not vary a great deal, and a slight effect of Reynolds number on Clq (Cl^ for 
ax = 0.0°) is observed in figure 5(a). The curve for Re = 2.76 x 10^ shows the classic stall break m 
the vicinity of ax = 12.0°. 



The drag polars of figure 5(b) show the differences in drag loads between the higher and lower 
values of Reynolds number. Many factors play a role in these results and in the drag plots that follow 
in this section. A better understanding can be gained by reviewing the drag buildup process in 
greater detail. 

The total tip drag for this type of configuration may be expressed in terms of its four main 
components: 

Cdt = Cdo + Cd^ = Co,„,, + Cd,^ + Co,, + Cd. (2) 

The levels of induced drag (Cq. ) are expected to be similar for points on different C^^ vs Cl^p 
curves with the same lift coefficient. The large difference between the curve for Re = 2.76 x 10^ and 
the other curves in figure 5(b) is probably related to the transition from laminar to turbulent flow. 
Reference 6 lists the flat-plate transition Reynolds number as somewhere between 3.5 x 10^ and 
10 X 105. For Re = 3.91 X 105 and above, the turbulent region in the boundary layers probably grows 
and thus increases the Cp^, component in equation (2). Also, turbulent boundary layers are gener- 
ally thicker than laminar ones (ref. 6); this allows greater interaction between the lifting surfaces and 
may increase the Cd.^^ term in equation (2). In particular, the increase in Cd;,, could occur when 
the tip caps of the two surfaces (facing each other across the small gap) are very close, such as for 
AO = -3.5°. 

Figure 6 shows the lift (fig. 6(a)) and drag (fig. 6(b)) trends of the RC10/08 tip for A0 = -13.0°. 
Figure 6(a) shows a loss of lift for Re = 2.76 x 10^, probably caused by partial flow separation, and 
subsequent reattachment, at a very low tip angle of attack. At this value of AG, when aj =0.0°, then 
aw = 13-0°. and the separation effect is related to the upwash of the inboard wing, which increases 
the loading on the tip surface. The separation and reattachment shown in figure 6(a) shows the high 
sensitivity of the flow to disturbances at low Reynolds numbers. 

Also, it is apparent from figure 6(a) that for a slightly higher Reynolds number 
(Re = 3.91 X 105), the flow is able to overcome adverse pressure gradients near the trailing edge, 
thus resulting in a more gradual loss of lift. The associated drag polar in figure 6(b) shows that in the 
range 0. 1 < Cl ^ 0.5 there is little Reynolds number effect. The drag data acquired for 

Re = 2.76 X 105 at this tip incidence angle were too unreliable to be shown here. 

Figures 7 and 8 show the lift curves and drag polars for the RC 10/05 tip at AS = -2.5° and 
AO = -10.0°, respectively. Again, the lift-curve slopes of figure 7(a) remain approximately constant 
over the range of Reynolds numbers shown, although the sensitivity of Cl^ to Reynolds number is 
more obvious than for the RC 10/08 tip. Separation is obsei-ved for the lowest Reynolds number at 
AO = -2.5°, with one data point indicating possible reattachment. The drag polars in figure 7(b) are 
grouped together and show little evidence of a trend caused by a change in Reynolds number. Fig- 
ure 8(a) shows the premature stall associated with the stronger upwash at AO = -10.0° and 
Re = 2.76 X 10^ (aj = 6.0°, aw = 16.0°). Also, as before, a Reynolds number of 3.91 x 10^ is high 
enough to delay separation. 



The drag polars of figure 8(b) (AG = -10.0°) show much higher sensitivity to Reynolds number 
than those of previous cases; for a constant value of Cl^ , as Re increases, Ci>j. also increases. This 
seems to occur for values of Cl from 0. 1 to 0.6. As explained above, the subtle contributions of 
skin friction versus interference drag cannot be extracted from these data. The spread in the curves of 
figure 8(b) are more consistent than the spread in the curves for the RClO/08 tip (figs. 5(b) and 6(b)); 
this may be related to local pressure gradients, which also have a strong influence on the location of 
the transition point. As will be discussed later, the wing upwash has a substantial impact on the tip- 
surface pressure distribution. Thus, these trends must be due to increases in Cd^^ and Cq.^^ . 



Tip Incidence Angle (AO) Effects 
RClO/08 Tip 

Some of the aerodynamic benefits derived from this configuration can be readily understood by 
comparing values of Cl„ and C^.^, for different values of AQ. Figures 9 and 10 show lift curves 
and drag polars for the RC 10/08 tip at Reynolds numbers of 4.79 x 10^ and 5.53 x 10^, respectively. 
For a given angle of attack, the lift produced by the tip increases as the negative value of AG 
increases. This is caused by the upwash effect of the wing vortex, which increases the local aerody- 
namic angle of attack across the span of the tip. 

For a given value of aj, the value of a^ actually increases with larger negative values of AG 
(ttw = -AG + aj). For large aw > the wing loading is high, resulting in a strong wingtip vortex 
(actually shed at the junction between wing and tip surfaces) and thus a high loading on the tip. For 
example, at aj = 0.0° and AG = -13.0°, then aw = 13.0° and figure 9(a) shows that the wing 
loading is very high. For the same ttj, as AG is reduced to -9.0° (ttw =9.0°) and then to -3.5° 
(aw = 3.5°), Cl is reduced accordingly. The plots only show points in the lower range of lift 
coefficient. This is because of concerns regarding allowable loads on the internal balance and its 
attachment point at high dynamic pressures. 

The drag plot (fig. 9(b)) shows that if a constant value of Cl.j. is considered, C^ decreases as 
AG increases negatively from 1.5° to -9.0°. This again is related to the reduction of induced drag that 
results from the stronger upwash associated with higher negative AG. It was shown through the use 
of a panel method code in reference 5 that this effect occurred consistently for negative AG. How- 
ever, because of the limitations of the numerical model, only induced drag was computed. The 
exception of figure 9(b) is that the drag curve for AG = -13.0° lies above that for AG = -9.0°, indi- 
cating a higher level of drag for the same tip lift. Although the differences are small, the drag polars 
of figure 10(b) (at the higher Reynolds number) also show that less drag was measured at 
AG = -3.5° than at AG = -13.0° for values of Cl.j. below 0.3. Although fewer data points are 
available at Re = 5.53 x 10^ and q = 20 psf, the curves of figure 10 show essentially the same 
trends as those of figure 9. 

Figure 1 1 shows the variation of tip pitching moment with lift for the RC 10/08 tip at both test 
conditions (Re = 4.79 x 10^ and 5.53 x 10^) considered above. In both cases, for a constant value of 
Cl , the pitching moment is less negative for larger negative AG. The reason for this is related to the 
strength of the wing vortex and its position relative to the tip surface. At negative AG, the tip trailing 



edge lies above the wing trailing edge, so the vortex shed from the wing/tip junction acts primarily 
on the lower surface of the tip. Since for constant Cl-p , cc w is larger for larger negative values of 
A6, the wing vortex strength and associated suction are also greater. Because the lower pressure 
region acts on the lower tip surface aft of the moment reference center, the resulting change in pitch- 
ing moment is in the nose-up direction, becoming less negative, as illustrated in figures 1 1(a) 
and (b). 

Figure 12 shows the variation of tip pitching moment with tip geometric angle of attack. For con- 
stant aj, the pitching moment is more negative for larger negative values of AG. This is primarily 
related to the increased lift at negative A6 illustrated in figures 9(a) and 10(a). 

RClO/05 Tip 

Figures 13 through 16 show the test results for the RClO/05 tip. As stated earlier, the only differ- 
ence between the two tip surfaces is the type of airfoil sections used and their spanwise distribution. 
The results are presented here for Reynolds numbers of 2.76 x 10^ and 3.91 x 10^ (different from 
those discussed in the previous section) to verify the observations made at the higher Reynolds num- 
bers. In figure 13(a) (Re = 2.76 x 10^), all three curves show some loss of lift because of flow sepa- 
ration. Note that for A9 = -8.0° and -2.5°, the flow reattaches, as indicated by the recoveries in lift 
beyond the initial separation point. Initial separation occurs in the same range of aj for 
AG = -10.0° and -8.0°, because the upwash on the tip does not differ much between these two cases. 
For a given value of aj, the upwash (and therefore the loading) on the tip is reduced for lower 
values of negative AG, so the angle of incidence of the entire configuration (and therefore aj) can 
be increased to a higher value before tip stall occurs, as seen in figure 13(a). 

The drag plot for Re = 2.76 x 10^ (fig. 13(b)) shows little difference between the AG = -10.0° 
and AG = -8.0° cases. Nonetheless, there is a large difference between the AG = -2.5° case and the 
AG = -10.0° and -8.0° cases, further supporting the conclusion of a reduction in induced drag at 
higher negative AG. The form drag for this tip is quite low (some of the airfoil sections are only 5% 
thick), and the total drag is reduced further by the drop in induced drag that results from the wing 
upwash. At optimum tip lift coefficients for AG = -8.0° and -10.0°, the tip drag loads were so small 
that they were below the sensitivity limit of the internal balance, hence the data points in the 
Cp =0.0 range of figure 13(b). 

Figure 14 shows lift and drag plots for Re = 3.91 x 10^. In figure 14(b), the data indicate a 
reduction in drag with increasing negative AG for the range -8.0° < AG < -2.5°. The curve for 
AG = -10.0° lies above that for AG = -8.0°. Note the similarity in the trends of figures 14(b) 
and 9(b). This may indicate that there is indeed a limit to which AG can be increased negatively and 
still achieve a reduction in total drag. This effect is related to the relative magnitudes of the four 
terms in equation (2). As stated earlier, a detailed drag breakdown cannot be established because of 
the limited scope of this test. 

The pitching moment curves of figures 15 and 16 show essentially the same behavior as 
observed in figures 1 1 and 12 for the RClO/08 tip. Slight differences in the slopes SC^.^, /3ax of 
figure 16 as AG changes can probably be attributed to experimental error. 



Flow Visualization 

The intent of the flow visualization runs was to study the surface-flow characteristics of the tip 
planforms, not to establish a direct relationship with the quantitative data discussed previously. How- 
ever, the quantitative data presented in previous sections are used as a guide in the interpretation of 
the flow visualization photographs. All results presented in this section were acquired at a dynamic 
pressure of 10 psf, with a corresponding Reynolds number of 3.91 x lO^. 

Figures 17(a) and (b) illustrate the upper and lower surface flow patterns on the RC 10/08 tip for 
ttj = -5.0° and AG = -13.0°. The data plotted in figure 6(a) show that at aj = -5.0° the tip is still 
generating positive lift, although at a slightly higher Reynolds number than for the present case. In 
the leading-edge region on both surfaces, the oil patterns confirm that the flow is still attached; how- 
ever, most of the flow is characterized by the formation of river patterns. As explained in refer- 
ence 7, this is caused by the excessive application of oil on the model surface. As seen in the pho- 
tographs, the influence of pressure gradients and shear stresses that normally determine the surface 
oil pattern is completely overcome by the effect of gravity on the thick oil film. 

For aj = -3.0° (fig. 18), the direction of the flow changes by roughly 90° on the upper surface 
in the vicinity of the junction between the wing and the tip. The surface flow is drawn toward the 
low-pressure region caused by the presence of the wing vortex; at this condition, aw = 10-0° and the 
wing loading is quite high. Clearly, the flow is still attached in this region. The change in direction of 
the surface flow is not observed on the lower surface (fig. 18(b)). As the wing angle of attack is 
increased at this constant value of AO, portions of the tip upper-surface flow continue to be drawn 
toward the wing and remain attached. At aj = -1.0° (fig. 19), despite the negative tip angle of 
attack, the loading is positive and moderately large (Cl.^ = 0.4, see fig. 6(a)). 

At ttj = 2.0° (fig. 20, ttw = 15.0°), a sharp separation line with reattachment is defined along 
the swept portion of the upper-surface leading edge. The tip inboard-surface flow is still being drawn 
toward the wing/tip juncture and is attached. This is of particular significance since it indicates that 
despite strong interactions between tip vortex and wing vortex, the tip surface flow does not separate 
prematurely in this area. The character of the flow on the lower surface did not change appreciably 
throughout the angle-of-attack sweep. 

The photographs for the RC 10/05 tip surface at AG = -2.5° are shown in figures 21 through 24. 
At ttx = -2.5° (fig. 21), river patterns caused by excessive oil are observed, as before. As aj is 
increased to 2.5° (fig. 22), a small region of separated flow develops on the upper surface close to 
the leading edge. Downstream of this area, the flow has reattached. The change in direction of the 
flow toward the wing/tip junction is also clearly discemable here. On the lower surface, the flow is 
attached and smooth well beyond the leading edge, but river patterns dominate the aft portion of the 
surface. 

At ax = 7.5° and 12.5° (figs. 23 and 24), the region of separated flow increases in size near the 
outboard end of the tip. This is an example of classic flow separation in the tip region of a swept sur- 
face. The accumulation of mass within the boundary layer, resulting from spanwise flow, alters the 
pressure gradient such that the flow initially separates in the outboard region. The lift curves shown 
in figures 13(a) and 14(a) corroborate the conclusion that the flow is separating at this condition. At 

8 



low Reynolds numbers, the plots show the onset of flow separation at about a^ = 8° to 9° for 
Ae = -2.5°. 



Tip Aerodynamic Parameters 

Tip Aerodynamic Center 

The location of the tip aerodynamic center along the mean aerodynamic chord is an important 
design parameter for free-tip planforms. For a tip oscillating about a mean angle of attack, the fre- 
quency of damped free vibration is directly related to the magnitude of the tip aerodynamic spring 
effect. This magnitude primarily depends on the ability of the tip to generate a restoring moment 
about the pitch axis. Since the resultant lift acts at the aerodynamic center, favorable restoring 
moments are obtained for aft aerodynamic center locations. 

The locations of the tip aerodynamic centers projected along the x axis of the wing-fixed coor- 
dinate system are calculated here from the data acquired for both tips. The goal is to determine 
whether AG, which is a function of wing angle of attack (and thus is influenced by wing vortex 
strength), has any impact on tip aerodynamic-center location. 

With the loads acting at the aerodynamic center, the tip pitching moment about an arbitrary point 
x along the wing chord may be expressed as 



C 



= C„,^ + (Cl, cos „,)(i^]. (c„, sin a,)(^) (3) 

where x is measured along the axis shown in flgure 2, with the positive direction pointing aft. The 
parameter x^c/cx is the projection on this axis of the tip aerodynamic center location. If equa- 
tion (3) is differentiated with respect to Cl.j. , the term dC^ /dCi^^ drops out by definition of the 
aerodynamic center. Also, if aj is assumed to be small and the contribution of the drag is 
neglected, the following expression is obtained: 



dCm-r 



X x-x 



ac 



dCLT- ct 



(4) 



In the present study, since the tip pitching moment was measured by a balance located at the 
inboard wing quarter-chord point, the value x/cy = 0.0 (see fig. 3) can be substituted into equa- 
tion (4), and the aerodynamic-center location can be easily calculated once the slope dCmQ q /^Cl^j, 
is known. The slopes for different values of Re and AO were calculated by fitting first-order curves 
to the test results presented earlier; these slopes are listed in tables C3 and C6 of appendix C. Fig- 
ure 25 shows the resulting locations of the aerodynamic center for the two tips, as a function of A9. 
For both tips, the position of the aerodynamic center is relatively constant with A6. For the RC 10/08 
tip, figure 25 shows variations with Reynolds number of up to 6% of the chord length. 



Lift and Pitching Moment Data 

The tip-planform lift and pitching moment slopes were computed from the complete body of 
experimental data and are listed in tables CI to C6 in appendix C. A quick check of the results shows 
no variation with A9 in the values of 3Cl^ fdaj and dC^^^ Ida-i, for a fixed value of Reynolds 
number. This is expected, since these parameters are typically affected only by planform geometry 
and Mach number. 

The y-intercepts for each curve are also listed in appendix C (tables CI to C6). These are the 



coefficients C t = C t 



„^.0°^"'^'^'"0"^'"T 



and C„ =C„ 



The first two 






parameters are plotted in figure 26 for the RC 10/08 tip. Generally, the value of Clq increases with 
negative AG. This is due to the strength of the upwash, which increases with ay/ • Figure 26(b) 
shows the associated effect on Cn,„ , which increases negatively with negative AG. This is due to 
the effect of increasing Cl , since it was shown above that the location of the tip aerodynamic 
center is essentially invariant with AG. 

Figure 27 shows the variation of Cl^ and C^j^ with AG for the RC 10/05 tip. The same trends 
as were observed in figure 26 are seen here, but with more consistency. The data seem to suggest a 
linear variation of these parameters with AG. The variations with Reynolds number are probably 
caused by boundary-layer thickness effects on chordwise pressure distributions. Figures 26 and 27 
thus quantify the variation with a^ of the upwash on C^^^ and C^^ caused by the strong influ- 
ence of aw on the AG parameter. 



USE OF STEADY FLOW EXPERIMENTAL DATA FOR CALCULATION OF FREE-TIP 

MOTION 

Experimental data such as those presented in this report can be used to bridge the gap between 
the cases of the deflected tip in steady flow and the free vibration in a free stream of an indepen- 
dently mounted tip surface pitching about the quarter-chord of the inboard wing. From an analytical 
standpoint, the free-oscillation case is a complex problem in which the governing equations of fluid 
dynamics must be solved simultaneously with the tip single-degree-of-freedom equation of motion. 
The equation of motion that appears in reference 8 for a two-dimensional rotary-wing aerodynamic 
environment is rewritten as follows for a tip mounted on a fixed wing of finite span: 

I„cxt +Ks(aT -Gp) = Cn,.^j^^pVl JStCt (5) 

This formulation applies only to a tip surface oscillating about an axis along the chord of a rigid 
wing, with a^ fixed. The term on the right-hand side is the aerodynamic moment and includes 
aerodynamic spring and damping effects. The evaluation of C^^ is the primary subject of this 
section. 



10 



The results presented in the current report focus on the variation of aerodynamic loads as a func- 
tion of A6. Although important information can be derived from data in this form, a different 
approach can be developed in which the parameters appearing in the expression for A6 are consid- 
ered independently, along with time-dependent variables. For the case of general unsteady motion of 
the tip surface with zero yaw (with a^ assumed fixed), the tip pitching-moment coefficient, vary- 
ing as a function of time, can be written as 

Cmj =CmT[«W. CCtW. hT(t),q.j.(t)] (6) 

For the case of arbitrary motion, two rigid-body degrees of freedom have been included: (1) wing 
plunging motion, which produces the h-p (t) term, and (2) the tip pitch rate, q j(t), which accounts 
for the time rate of change of tip incidence angle relative to a horizontal reference plane. A Taylor 
series expansion of Cj^ about the point where a^ =aT = hx=q^ =0 can be written as 
follows: 

C^.j,(aw,t) = C^Q -h(aCjn.j,/3aw)aw +(5CmT./aaT)aT 

+ [3C^.^/a(hT/V.)](hT/V<.) + [8Cn,^/a(qTCT/Voo)](qT,CT/V^) (7) 

This expansion is linearized by neglecting second- and higher-order terms, since the angles and 
rates are assumed to be small. The partial derivatives are evaluated at the point about which the 
series is expanded. The second and third terms in the series will be discussed first. 

Equation (7) cannot be used with data acquired in terms of A6, since 3/3(A6) terms (in an 
expansion with A0 as the independent variable) imply a variation of both ttw and ttj simultane- 
ously, thus violating the definition of the partial derivative. To determine the first two slopes of the 
Taylor series expansion, the data would have to be acquired in a different manner; instead of varying 
the angle of incidence of the entire configuration with the tip deflected at a fixed angle relative to the 
wing (A9), the measurements would have to be taken at fixed values of a^ while the tip angle of 
attack was varied. Thus, families of curves of C^,^ vs a^ would be obtained for constant values of 
ttw ■ The derivative dC^^ /8aw would then easily be obtained by measuring the change in C^ ^ 
among the different curves, for aj = 0.0°. The second partial derivative is simply the slope 
3Cm-p /^OLj of the tip pitching moment coefficient curve for a^ = 0.0°. Unfortunately, this 
approach was not implemented for the present study because the test focused on deflected tips, for 
which data have traditionally been presented as a function of A0 (refs. 1-5). 

The last two terms in the series expansion (eq. (7)) follow from a quasi-steady formulation in 
which CJ^^(a^f^, t) depends on the instantaneous values of h j (t) and q^(t). One problem with this 
approach is that it is based on the assumption that dC,^ /dihjY^j and dCj^^ /d^q^Cj/W^j are 
constant and can be evaluated or measured experimentally. A more serious problem with 
equation (7) is that it does not acknowledge any effects that the past history of the motion may have 
on the instantaneous aerodynamic response. 



11 



Following the methods of indicial unsteady aerodynamic theory presented in reference 9, the 
pitching moment coefficient may be formulated to include the effects of previous motion on the cur- 
rent loading, as follows: 

C^^ (aw, t) = C^^ + (ac^^ /3cxw)«w + j^C^^ l^ " ''' «t(t^)' ^1(1^)] ^^^^t 

+ ^{c,Jt-T;aT(x),qT(t)]^dT (8) 

The first parameter in the integrands represents the incremental pitching moment response at 
time t that corresponds to a unit step change in aj and q^ that occurs at time x. If ajCx) and 
q^(T) are approximated by a superposition of unit step changes, then the pitching moment response 
is simply the summation of incremental responses to unit step changes in aj(x) and <{j(x). Equa- 
tion (8) has already been simplified in that C^^ [. . .] and C^q [• • .] are based only on the values of 
ax and qj at the origin of the step. In principle, the responses should be based on the complete his- 
tories of ttj and q^ . The histories could be reconstructed by Taylor series expansions about the 
time of the step initiatiqn (for example, aj(t') = aj(T) - dj(x)(T- f) - ..., where f represents a 
time earlier than x); however, as explained in reference 9, the indicial responses should have 
"forgotten" long-past events. Also, limiting the present analysis to that of slowly varying motions 
allows for the exclusion of djandq.j, terms. 

A final assumption is required before a working form of equation (8) can be obtained. If aj(t) 
and GxCt) are represented as harmonic functions, powers of q.^. (t) higher than the first will be of 
second and higher orders in frequency [q.j, (t) = Qjit)] After expressing the integrands of 
equation (8) in terms of deficiency functions and expanding these in Taylor series about q.^ (0) = 0, 
we may neglect all terms containing powers of q.j.(t) higher than the first, since the frequency is 
assumed to be small. 

The expression that results from integrating equation (8) contains the term Cj^^ [=»; aj(t), 0] 
described in reference 9 as the pitching moment coefficient measured in a steady flow with a^ 
fixed, at the instantaneous value aj(t) = and q.^ (t) = 0. The appearance of this term is the key to 
the use of experimental data in the computation of free-tip oscillatory motion. In the context of the 
rigid semispan wing with free-pitching tip, this term takes the following form: 

Cmxh; «T(t)> 0] = C^o +(dC^j/da^^)aw+(^C^jftaj)aj{t) (9) 

The first three coefficients are parameters that can be evaluated experimentally, as discussed 
previously. The integrated version of equation (8) can be further simplified by the substitution 
dj = q.j., which holds for rectilinear motion in the present configuration. 

Substituting equation (9) into the simplified version of equation (8) leads to the final working 
form, which can be used for free-tip analysis and is given by 



12 



C,^(aw,t) = C^o +(^C,T/3«w)aw +(ac.T/a«T)aT(t)H- {Cn. JaT(t)]} ^^^ 

The last term is analogous to the pitch-damping stability derivative. In this case however, it must 
be evaluated or measured for each value of ajCt) within the expected range of the oscillatory 
motion. This may be done by prescribing a small-amplitude harmonic oscillation using each 
expected value of ajCt) as a mean. The required data can be obtained from experimental measure- 
ments or by solving numerically the appropriate unsteady field equations of fluid mechanics, for 
pure harmonic pitching motion. For a more comprehensive data base of coefficients C^^ [aj(t)J, 
the prescribed motion can be repeated for several frequencies. The most efficient and cosT-effective 
approach to achieve this for the free-tip on a semispan wing would be to use a panel-method com- 
puter code for unsteady motion, such as the one described in reference 10. The coefficient is evalu- 
ated from the component of the pitching-moment-response time history that is 90° out of phase with 
the prescribed angle of attack. This approach was used successfully in reference 1 1 for the case of an 
oscillating flap on an airfoil in transonic flow. For that study, the time-dependent Euler equations 
were solved to obtain the desired coefficients. 

The advantage of this approach is obvious. Once the coefficients for a given tip configuration 
have been tabulated or curve-fitted, the oscillatory motion of a free-pitching tip may be calculated 
from equation (5) by a simple numerical integration scheme, and by obtaining C^^ (« wO ^^°"^ 
equation (10) with current values of aj{t) and djCt). Hence, the mechanical characteristics of the 
tip (I„, Kg) can be altered and the tip motion recomputed as long as the expected response frequency 
is within the range of prescribed-motion frequencies used to evaluate C ^^ [a j (t)J . 

The preceding discussion provides a framework for future research in the analysis of free-pitch- 
ing tip motion. The relationship between the evaluation or measurement of steady aerodynamic loads 
and the solution of the unsteady tip-motion problem has been cleariy established. 

CONCLUDING REMARKS 

The aerodynamic characteristics of two new tip surfaces proposed for a small-scale Free-Tip 
Rotor model were measured at dynamic pressures from 5 to 25 psf and Reynolds numbers from 
2.76 X 105 to 6.18 X lO^. The tips had similar planform characteristics, but different airfoil distribu- 
tions, varying from thickness ratios of 10% on the inboard portion to 8% or 5% toward the outboard 
edge of the tip. The major issues addressed in this study are as follows: 

1 . Lift and pitching moment data support the theory that upwash from the semispan wing has a 
strong influence on the aerodynamic loading of the deflected tip. This effect is proportional to the 
wing angle of attack. 

2. The upwash from the inboard wing reduces the stall angle of the tip surface, especially at 
lower Reynolds numbers. A slight increase in Reynolds number is sufficient to delay this effect. 



13 



3. Experimental evidence suggests that the reduction in tip induced drag associated with the 
wing upwash reduces the total tip drag. However, there is apparently a negative value of AO for 
which the drag reduction reaches its limit. 

4. The flow visualization results show no outstanding features that might be detrimental to the 
performance of the tips. As the lift coefficient is increased, the flow separates gradually on the out- 
board portion of the tip. Surface flows on the inboard portion of the tip are drawn toward the gap 
between the wing and the tip, but remain attached up to the inboard side edge of the surface. 

5. Lift and pitching-moment slopes, as well as coefficients for zero angle of attack, have been 
computed for both tips. The test results indicate that the surface aerodynamic centers are sufficiently 
aft of the quarter-chord pitch axis to be suitable for free-tip designs. The variations in aerodynamic 
center position of the tip are minimal and are not believed to be due to changes in wing loading. 

6. A new approach that treats wing and tip angles of attack as independent parameters was pro- 
posed. The required Taylor series coefficients can be easily obtained from test results by graphical 
means. These can, in turn, be combined with tabulated values of tip pitch-damping coefficients to 
obtain the time-varying pitching moment response of a free-tip. This method allows the calculation 
of general unsteady tip motion without repeatedly solving the coupled single-degree-of-freedom/ 
aerodynamic problem. 



14 



APPENDIX A 
AIRFOIL SURFACE ORDINATES AND WING TWIST DISTRIBUTION 

This appendix contains the following tables and figures: 
Table Al. Surface coordinates for the RC05 airfoil section. 
Table A2. Surface coordinates for the RC08 airfoil section. 
Table A3. Surface coordinates for the RCIO airfoil section. 
Table A4. Surface coordinates for the V23010 airfoil section. 
Figure Al. Semispan wing-twist distribution. 



15 



Table Al. Surface coordinates for the RC05 airfoil section 



x/c* 


z^/c** 


z,/ct 


x/c* 


z„/c** 


Zj/Ct 


0.000000 


0.000000 


0.000000 


0.444581 


0.027785 


-0.022323 


0.001007 


0.003369 


-0.002282 


0.476209 


0.027498 


-0.022286 


0.004023 


0.006872 


0.004436 


0.507933 


0.027046 


-0.022115 


0.009036 


0.010424 


-0.006447 


0.539625 


0.026399 


-0.021793 


0.016026 


0.013926 


-0.008314 


0.571157 


0.025542 


-0.021316 


0.024964 


0.017264 


-0.010033 


0.602403 


0.024481 


-0.020683 


0.030154 


0.018834 


-0.010844 


0.633237 


0.023223 


-0.019893 


0.041946 


0.021703 


-0.012369 


0.663534 


0.021805 


-0.018951 


0.063075 


0.025143 


-0.014439 


0.707707 


0.019486 


-0.017251 


0.071008 


0.026028 


-0.015074 


0.736136 


0.017902 


-0.015931 


0.088162 


0.027389 


-0.016253 


0.750000 


0.017120 


-0.015216 


0.097365 


0.027870 


-0.016797 


0.763613 


0.016350 


-0.014471 


0.116978 


0.028482 


-0.017791 


0.790028 


0.014850 


-0.012895 


0.138133 


0.028707 


-0.018650 


0.827430 


0.012667 


-0.010414 


0.160745 


0.028667 


-0.019374 


0.850737 


0.011192 


-0.008760 


0.172570 


0.028586 


-0.019687 


0.872632 


0.009658 


-0.007178 


0.197195 


0.028371 


-0.020221 


0.902635 


0.007258 


-0.005058 


0.236387 


0.028062 


-0.020833 


0.928992 


0.004954 


-0.003354 


0.250000 


0.027990 


-0.020997 


0.951463 


0.003090 


-0.002104 


0.277967 


0.027907 


-0.021289 


0.975036 


0.001695 


-0.001079 


0.292293 


0.027892 


-0.021422 


0.983974 


0.001472 


-0.000784 


0.321557 


0.027905 


-0.021671 


0.995977 


0.001658 


-0.000515 


0.351540 


0.027946 


-0.021900 


0.997736 


0.001756 


-0.000506 


0.382121 


0.027967 


-0.022102 


0.999748 


0.001897 


-0.000547 


0.413176 


0.027928 


-0.022252 


1.000000 


0.001900 


-0.000600 


*x/c is the chordwise airfoil station for surface definition. 




**zjc is the airfoil upr 


)er-surface ordinate. 







' zj/c is the airfoil lower-surface ordinate. 



16 



Table A2. Surface coordinates for the RC08 airfoil section 



x/c* 


z„/c** 


zi/ct 


0.000000 


0.000000 


0.000000 


0.003140 


0.006710 


-0.006560 


0.011700 


0.013120 


-0.010960 


0.025340 


0.019000 


-0.014450 


0.043690 


0.024550 


-0.017290 


0.065080 


0.029520 


-0.019510 


0.088890 


0.033890 


-0.021290 


0.114710 


0.037700 


-0.022780 


0.142390 


0.040960 


-0.024030 


0.171860 


0.043760 


-0.025130 


0.202980 


0.046060 


-0.026090 


0.235600 


0.047890 


-0.026930 


0.269590 


0.049280 


-0.027910 


0.304840 


0.050230 


-0.028530 


0.341210 


0.050750 


-0.028770 


0.378400 


0.050870 


-0.029130 


0.416120 


0.050570 


-0.029340 


0.454040 


0.049860 


-0.029390 


0.491990 


0.048760 


-0.029290 


0.529830 


0.047250 


-0.029020 


0.567660 


0.045330 


-0.028580 


0.605530 


0.043000 


-0.027960 


0.643670 


0.040240 


-0.027130 


0.682380 


0.037040 


-0.026050 


0.721500 


0.033420 


-0.024690 


0.760570 


0.029440 


-0.023000 


0.798420 


0.025280 


-0.020990 


0.834570 


0.021070 


-0.018670 


0.869270 


0.016860 


-0.015990 


0.902800 


0.012710 


-0.012930 


0.935430 


0.008640 


-0.009420 


0.967820 


0.004900 


-0.005130 


1.000000 


0.001300 


-0.000500 



*x/c is the chordwise airfoil station for surface 
definition. 



** 



Zy/c is the airfoil upper-surface ordinate, 
tzj/c is the airfoil lower-surface ordinate. 



17 



Table A3. Surface coordinates for the RCIO airfoil section 



x/c* 


Zu/c** 


Zj/Ct 


0.000000 


0.000000 


0.000000 


0.003100 


0.009060 


-0.007650 


0.010930 


0.017000 


-0.012300 


0.024030 


0.024620 


-0.016190 


0.042140 


0.031930 


-0.019500 


0.063470 


0.038540 


-0.022030 


0.087240 


0.044350 


-0.024000 


0.113050 


0.049440 


-0.025610 


0.140750 


0.053760 


-0.026960 


0.170230 


0.057480 


-0.028170 


0.201370 


0.060510 


-0.029230 


0.234020 


0.062920 


-0.030200 


0.268070 


0.064740 


-0.031070 


0.303430 


0.065960 


-0.031840 


0.339950 


0.066610 


-0.032490 


0.377330 


0.066720 


-0.033010 


0.415240 


0.066270 


-0.033370 


0.453360 


0.065580 


-0.033570 


0.491480 


0.063760 


-0.033600 


0.529490 


0.06i700 


-0.033460 


0.567460 


0.059090 


-0.033130 


0.605470 


0.055930 


-0.032600 


0.643780 


0.052200 


-0.031840 


0.682760 


0.047870 


-0.030790 


0.722320 


0.042960 


-0.029390 


0.761850 


0.037580 


-0.027580 


0.800020 


0.032000 


-0.025360 


0.836230 


0.026440 


-0.022720 


0.870820 


0.020960 


-O.019590 


0.904090 


0.015640 


-0.015940 


0.936310 


0.010520 


-0.011710 


0.968280 


0.006000 


-0.006330 


1.000000 


0.001800 


-0.000200 


*x/c is the chordwise airfoil station for surface 


definition. 







'*Zu/c is the airfoil upper-surface ordinate, 
tzj/c is the airfoil lower-surface ordinate. 



18 



Table A4. Surface coordinates for the V23010 airfoil section 



x/c* 


Zu/C** 


zi/ct 


0.00000 


-O.02250 


-0.02250 


0.00500 


-0.00780 


-0.03290 


0.01000 


-0.00240 


-0.03620 


0.01500 


0.00190 


-0.03780 


0.02500 


0.00960 


-0.03940 


0.03500 


0.01550 


-0.04040 


0.04700 


0.02140 


-0.04120 


0.06000 


0.02650 


-0.04200 


0.08000 


0.03270 


-0.04340 


0.11000 


0.03960 


-0.04490 


0.15000 


0.04550 


-0.04710 


0.19000 


0.04890 


-0.04940 


0.23000 


0.04990 


-0.05130 


0.27000 


0.04990 


-0.05220 


0.31000 


0.04970 


-0.05215 


0.35000 


0.04900 


-0.05170 


0.39000 


0.04800 


-0.05050 


0.43000 


0.04650 


-0.04870 


0.47000 


0.04460 


-0.04680 


0.51000 


0.04240 


0.04400 


0.55000 


0.03970 


-0.04120 


0.59000 


0.03690 


-0.03800 


0.63000 


0.03360 


-0.03460 


0.67000 


0.03010 


-0.03080 


0.71000 


0.02630 


-0.02690 


0.75000 


0.02230 


-0.02260 


0.79000 


0.01810 


-0.01820 


0.83000 


0.01370 


-0.01360 


0.87000 


0.00930 


-0.00930 


0.91000 


0.00560 


-0.00570 


0.94500 


0.00280 


-0.00310 


0.96000 


0.00235 


-0.00235 


1.00000 


0.00235 


-0.00235 


*x/c is the chordwise airfoil station for surface 


definition. 






**zjc is the airfoil upper-surface ordinate. 



tzj/c is the airfoil lower-surface ordinate. 



19 




Tunnel 
floor 



Figure Al. Semispan wing-twist distribution. 



20 



APPENDIX B 

WIND TUNNEL CORRECTIONS 

Standard wind tunnel wall corrections were applied in accordance with the methods outlined in 
reference 12. As recommended by the authors for such cases, the semispan wing was modeled as a 
full-span surface with twice the length of the half-span surface. The fictitious wind tunnel was also 
assumed to have a ratio of breadth over height (B/H) of 14 ft/10 ft = 1.40. The following is a sum- 
mary of these corrections: 

Velocity Correction for Solid-Body Blockage and Wake Blockage 

The total velocity increment is 

^ ~ ^sb "^ ^wb 

where 

Egb = f (wing volume, tunnel area, model span, t/c, B/H) 

e^b = f (wing area, tunnel area, wing drag coefficient) 
Since wing loads were not measured, an empirical estimate of wing drag was used. 

The corrected velocity is 



the corrected Reynolds number is 



and the corrected dynamic pressure is 



V^=Vu(l + e) 



Re = Reu(H-e) 



q = qu(l + 2e) 



where the subscript u refers to the uncorrected value of the parameter. For this test, e = 3.01 x lO"^ 
was computed. 



21 



Corrections for Downwash and Streamline Curvature 

During the data reduction process, corrections were applied to the following tip load parameters 
in the given order: 

The first correction was 

CL=CLjl-2e) + (ACL),, 

where 

(ACL) = -CL„(^a)sc 
(Aa)sc=t2(Aa)dw 

t2 = f (chord length, H/B, B) 
(Aa)dw = f (^L , equivalent vortex span, tunnel area, tip planform area, model chord length) 

For this test, the computed correction parameters yield 

Cl =Cl„ (1-2^)- (5-54 x10-5)Cl^ 

The second correction was 

ax=aT +(Aa)j^(l + t2) 



Therefore, 



aT,=aT +0.0373Cj (1 + 0.033) 



The third correction was 



CD=CD,(l-2e) + ACD 



where 

ACd = f (Cl. vortex span, tunnel area, tip planform area, model chord length) 



22 



Therefore, 

Cj3,=Cd (l-2e) + (6.51xl0^)cj^ 

The fourth correction was 

Cm=C (l-2e)-0.25(ACL) 



Therefore, 



'sc 



Cn.=C„, (l-2e)-0.25(-5.54xl0-5)C, 



23 



APPENDIX C 

TIP AERODYNAMIC PARAMETERS 

This appendix contains the following tables: 

Table CI. Lift-curve slopes and zero-alpha lift coefficients for the RC 10/08 tip. 

Table C2. Pitching-moment-curve slopes and zero-alpha pitching moment coefficients for the 
RC 10/08 tip. 

Table C3. Pitching-moment-curve slopes and zero-lift pitching moment coefficients for the 
RC 10/08 tip. 

Table C4. Lift-curve slopes and zero-alpha lift coefficients for the RC 10/05 tip. 

Table C5. Pitching-moment-curve slopes and zero-alpha pitching moment coefficients for the 
RC 10/05 tip. 

Table C6. Pitching-moment-curve slopes and zero-lift pitching moment coefficients for the 
RC 10/05 tip. 



PRECEDING PAGE BLANK NOT FfLMED 

25 



Table CI. Lift-curve slopes and zero-alpha lift coefficients for the 
RC 10/08 tip 



Re A0 Cl„ Clo 



(deg) (deg-Q 



2.76 X 105 19:0 0.0735 0.6308 

-3.5 0.0758 0.0865 

-3.5 0.0696 0.1100 

1.5 0.0765 0.1601 

3.91 X 105 _i3.o 0.0848 0.5067 

-9.0 0.0928 0.6214 

-3.5 0.0808 0.1060 

-3.5 0.0738 0.1055 

1.5 0.0805 0.0814 

4 79x105 _i3.o 0.0845 0.5031 

-9.0 0.0847 0.4469 

-3.5 0.0860 0.1214 

-3.5 0.0809 0.1142 

1.5 0.0845 0.0165 

5.53x105 -13.0 0.0875 0.5000 

-9.0 0.0866 0.4531 

-3.5 0.0822 0.1338 

1.5 0.0774 0.0275 

6.18x105 -13.0 0.0797 0.4733 

-3.5 0.0805 0.1302 



26 



Table C2. Pitching-moment-curve slopes and zero-alpha pitching 
moment coefficients for the RC 10/08 tip 



2.76 X 105 



3.91 X 105 



4.79 X 105 



5.53 X 105 



6.18 X 105 



Re AB C^^ C^Q 

(deg) (deg-i) 



-13.0 


-0.0216 


-0.1349 


-9.0 


-0.0158 


-0.1225 


-3.5 


-0.0197 


-0.0457 


-3.5 


-0.0198 


-0.0525 


1.5 


-0.0175 


-0.0536 


-13.0 


-0.0167 


-0.1254 


-9.0 


-0.0172 


-0.1139 


-3.5 


-0.0191 


-0.0468 


-3.5 


-0.0187 


-0.0542 


1.5 


-0.0161 


-0.0521 


-13.0 


-0.0230 


-0.1314 


-9.0 


-0.0212 


-0.1128 


-3.5 


-0.0222 


-0.0443 


1.5 


-0.0192 


-0.0319 


-13.0 


-0.0242 


-0.1308 


-3.5 


-0.0218 


-0.0444 


1.5 


-0.0196 


-0.0277 


-13.0 


-0.0219 


-0.1223 


-9.0 


-0.0238 


-0.1151 


-3.5 


-0.0215 


-0.0422 



27 



Table C3. Pitching-moment-curve slopes and zero-lift pitching moment 
coefficients for the RC 10/08 tip 



Re 


Ae 


dC^.j./dCL^ 


Cmo 




(deg) 


(deg-1) 




2.76 X 10^ 


-13.0 


-0.2529 


-0.0074 




-9.0 


-0.2090 


-0.0091 




-3.5 


-0.2410 


-0.0270 




-3.5 


-0.2624 


-0.0261 




1.5 


-0.2193 


-0.0242 


3.91 X 105 


-13.0 


-0.2382 


-0.0102 




-9.0 


-0.1865 


0.0021 




-3.5 


-0.2373 


-0.0214 




-3.5 


-0.2539 


-0.0273 




1.5 


-0.2122 


-0.0298 


4.79 X 105 


-13.0 


-0.2738 


0.0062 




-9.0 


-0.2437 


0.0044 




-3.5 


-0.2585 


-0.0129 




-3.5 


-0.2828 


-0.0196 




1.5 


-0.2266 


-0.0283 


5.53 X 105 


-13.0 


-0.2771 


0.0076 




-9.0 


-0.2681 


0.0070 




-3.5 


-0.2653 


-0.0088 




1.5 


-0.2526 


-0.0208 


6.18 X 105 


-13.0 


-0.2740 


0.0074 




-9.0 


-0.2435 


0.0131 




-3.5 


-0.2672 


-0.0074 



28 



Table C4. Lift-curve slopes and zero-alpha lift coefficients for the 
RC 10/05 tip 



Re 


AG 


Ct 


Clo 


2.76 X 10^ 


(deg) 
-10.0 
-8.0 

-2.5 


(deg-1) 
0.0687 
0.0600 
0.0763 


0.4344 
0.3449 
0.0901 


3.91 X 105 


-10.0 
-8.0 
-2.5 


0.0792 
0.0788 
0.0855 


0.4205 
0.2841 
0.0957 


4.79 X 105 


-10.0 
-8.0 
-2.5 


0.0808 
0.0802 
0.0832 


0.4043 
0.2982 
0.1747 


5.53 X 105 


-10.0 
-8.0 

-2.5 


0.0804 
0.0807 
0.0793 


0.3936 
0.2914 
0.1616 


6.18 X 105 


-10.0 

-2.5 


0.0848 
0.0760 


0.3955 
0.1498 



29 



Table C5. Pitching-moment-curve slopes and zero-alpha pitching 
moment coefficients for the RC 10/05 tip 



Re 


AB 
(deg) 


(deg-I) 


c 


2.76 xlO^ 


-10.0 
-8.0 
-2.5 


-0.0153 
-0.0141 
-0.0172 


-0.1002 
-0.0797 
-0.0331 


3.9 X 105 


-10.0 

-10.0 

-8.0 

-2.5 


-0.0150 
-0.0151 
-0.0155 
-0.0163 


-0.0922 
-0.0877 
-0.0742 
-0.0346 


4.79x105 


-10.0 

-10.0 

-8.0 

-2.5 


-0.0163 
-0.0153 
-0.0163 
-0.0171 


-0.0831 
-0.0852 
-0.0719 
-0.0340 


5.53 X 105 


-10.0 

-10.0 

-8.0 

-2.5 


-0.0159 
-0.0155 
-0.0159 
-0.0166 


-0.0779 
-0.0842 
-0.0675 
-0.0313 


6.18 X 105 


-10.0 
-10.0 

-2.5 


-0.0164 
-0.0159 
-0.0163 


-0.0753 
-0.0828 
-0.0288 



30 



Table C6. Pitching-moment-curve slopes and zero-lift pitching moment 
coefficients for the RC 10/05 tip 



Re 


AB 

(deg) 


(deg-1) 


Cmo 


2.76 X 10^ 


-10.0 
-8.0 

-2.5 


-0.2206 
-0.2369 
-0.2098 


-0.0043 

0.0017 

-0.0169 


3.9 X 105 


-10.0 

-10.0 

-8.0 

-2.5 


-0.2025 
-0.2017 
-0.2028 
-0.2088 


-0.0075 

0.0015 

-0.0164 

-0.0118 


4.79 X 105 


-10.0 

-10.0 

-8.0 

-2.5 


-0.2020 
-0.1959 
-0.2030 
-0.2047 


-0.0014 
0.0073 

-0.0114 
0.0017 


5.53 X 105 


-10.0 

-10.0 

-8.0 

-2.5 


-0.1981 
-0.1761 
-0.1967 
-0.2094 


0.0001 

0.0065 

-0.0102 

0.0026 


6.18 X 105 


-10.0 
-10.0 

-2.5 


-0.1930 
-0.1934 
-0.2141 


0.0010 
0.0249 
0.0033 



31 



REFERENCES 

1 . Stroub, R. H.; Young, L. A.; Keys, C. N.; and Cawthorne M. H.: Free-Tip Rotor Wind Tunnel 
Test Results. AHS J., vol. 31, no. 3, July 1986, pp. 19-26. 

2 Van Aken, J. M.: An Investigation of Tip Planform Influence on the Aerodynamic Load Char- 
acteristics of a Semispan Wing and Wing Tip. Report 5171-1, The University of Kansas 
Center for Research, Inc., Lawrence, KS, Dec. 1985. 

3. Van Aken, J. M.; and Stroub, R. H.: Tip Aerodynamics from Wind Tunnel Test of Semispan 
Wing. NASA TM-88253, 1986. 

4 Van Aken, J. M. : Experimental Investigation of the Influence of Tip Planform and Wing 

Sweep on the Tip Aerodynamic Load Characteristics. Report 7440-4, The University of 
Kansas Center for Research, Inc., Lawrence, KS, Nov. 1988. 

5. Martin, D. M. ; and Fortin P. E.: VSAERO Analysis of Tip Planforms for the Free-Tip Rotor. 

NASA CR-177487, 1988. 

6. Schlichting, H.: Boundary Layer Theory. McGraw-Hill, New York, 1968. 

7. Maltby, R. L.: Flow Visualization in Wind Tunnels Using Indicators. AGARDograph 70, Apr. 

1962. 

8. Yates, L.; and Kumagai, H.: Application of Two-Dimensional Unsteady Aerodynamics to a 

Free-Tip Rotor Response Analysis. NASA CR-177348, 1985. 

9. Tobak, M.; and Schiff, L. B.: Aerodynamic Mathematical Modeling— Basic Concepts. 

AGARD Lecture Series on Dynamic Stability Parameters. AGARD-LS-1 14, May 1981. 

10. Katz, J.; and Maskew, B.: Unsteady Low-speed Aerodynamic Model for Complete Aircraft 

Configurations. J. Aircraft, vol. 24, Apr. 1988. 

11. Chyu, W. J.; and Schiff, L. B.: Nonlinear Aerodynamic Modeling of Flap Oscillations in 

Transonic Flow: A Numerical Validation. AIAA J., vol. 21, no. 1, Jan. 1983. 

12. Pope, A.; and Harper, J.: Low-Speed Wind Tunnel Testing. Wiley, New York, 1966. 



PRECEDING PAGE BLANK NOT FILMED 

33 



Transition 
region 




All dimensions in ft 

Figure 1 . Geometry of tip planforms. 



34 



Deflected tip 
(metric section) 



Wing coordinate 
axis system 




Balance sleeve 
assembly 

Internal 
balance 



Balsa wood Insert 
(nonmetric) 



Steel wing 



Figure 2. Wing and tip layout with wing-fixed axis system. 



35 



DRIGINAL PAGE 
BLACK AND WHITE PHOTOGRAPH 




(a) Installation 
Figure 3. Semispan wing with deflected tip, in wind tunnel. 



36 



ORIGsNAL' PAOE 
BLACK AND WHITE PHOTOGRAPH 




(b) Close-up view 
Figure 3. Concluded. 



37 



Chord reference 
li 




N : jalanceriornial-force axis 
STBai ance side-force axis 
RM:^lance rolling-moment axis 



Figure 4. Orientation of internal-balance axis system and V23010 airfoil reference line. 



1.0 



.8 - 



.6 



-.2 



n 


Re=2.76X105 


o 


Re=3.91X10^ 


A 


Re=4.79X105 □ a 


+ 

X 


Re=5.53X105 □ ° 
Re=6.18X105 ^ ^ 




■/ 


- 


d/^ 




/ 




V 


- 


/ 


(a) 


1 1 1 J- 1 



4 8 

a^, (deg) 



12 



.10 r 



.08 



.06 



.04 



.02 



16 




Figure 5. Variation of (a) lift and (b) drag coefficients for the RC 10/08 tip, AS = -3.5°. 



38 



1.0 r 



1- 
-^ 4 



D Re=2.76X10=' 

O Re=3.91X10^ 

A Re=4.79X10^ 

+ Re=5.53X10^ 

X Re=6.18X10^ 




(a) 



-.2' 



4 8 

a,,, (deg) 



12 



O 

o 



.10 



.08 



.06 



.04 



.02 



I 

16 



oL 

-.2 



(b) 



,-ati - 



.2 



Figure 6. Variation of (a) lift and (b) drag coefficients for the RC 10/08 tip, AO 



= -13.0° 



1.0 



1.0 r 



.6 



o" ■* 



D Re=2.76X10° 
O Re=3.91X10^ 
A Re=4.79X10^ 
+ Re=5.53X10^ 
X Re=6.18X10^ 



(a) 




a^(deg) 



12 



.10 r 



.08 



o 



.06 



.04 



.02 



(b) 



16 



.2 




Figure 7. Variation of (a) lift and (b) drag coefficients for the RClO/05 tip, AO = -2.5' 



1.0 



39 



1.0 



-.2 



D Re=2.76X10^ 
O Re=3.91X105 
A Re=4.79X10^ 
+ Re=5.53X10^ 
X Re=6.18X10^ 




.10 



.08 - 



.06 - 



I- 

o 



.04 



.02 



(a) 



D d 



12 



16 




a,,,(deg) 



Figure 8. Variation of (a) lift and (b) drag coefficients for the RC 10/05 tip, AS = -10.0° 



1.0 



.8 - 



.6 - 



r ^ 


A0=-13.O° 




o 


A0=-9.O° 




A 


A0=-3.5'' 




+ 


Ae^l.B" 




_ 


P 






/ 


//• 




// 




- 


/ • 
/ 




(a) 


1 


J 1 1 1 



.10 r 



.08 



.06 - 



I- 

D 

o 



.04 



.02 



(b) .,..4- 
a 1 




4 8 

a^p (deg) 



12 



16 



.4 .6 



.8 1.0 



Figure 9. Variation of (a) lift and (b) drag coefficients for the RClO/08 tip, Re = 4.79 x 10^. 



40 



1.0 



^ .4h 



.2 - 



-.2 



n AS =-13.0° 
A Afl=-3.5° 
+ Ae=1.5° 




(a) 



/ 



.10 



.08 



.06 



I- 
Q 

o 



.04 



.02 



12 



16 



(b) 



a,p (deg) 




.2 



.6 



1.0 



Figure 10. Variation of (a) lift and (b) drag coefficients for the RC 10/08 tip, Re = 5.53 x 10^. 



.05 



- 



-.05 



e" -.10 



-.15 



.20 



-.25 



O Ae=-13.0'' 

O A0=-9.O° 

A Ae=-3.5" 

+ A9 = 1.5'' 



(a) 




□ A6l=-13.0'' 
A A61=-3.5'' 



(b) 




1.0 



-.2 



.4 



1.0 



Figure 11. Variation of pitching moment coefficient with lift coefficient for the RC 10/08 tip; 
(a) Re = 4.79 x 105, (b) Re = 5.53 x 10^. 



41 



.05 



.05 



E -.10 



-.15 



-.20 - 



D Afl=-13.0° 

O A0=-9.0° 

A Ae=-3.5° 

+ Ae = 1.5° 




(a) 
-.25 I L. 



D Ae=-13.0° 
A Ae=-3.5° 
+ A0 = 1.5'' 




"A 



(b) 



12 



16 



12 



16 



a,|,(deg) 



a^ (deg) 



Figure 12. Variation of pitching moment coefficient with angle of attack for the RClO/08 tip; 
(a) Re = 4.79 x lO^, (b) Re = 5.53 x lO^. 





l.U 


' 


D 


Ae= 


-10.0° 









A 


ts-- 


-8.0° , ^ 
-2.5° 4 » 




.8 








f 9 ." 




.6 












.4 


• 






?■'' / 

8'' / 




.2 








4' 

/4 





1 


(a) 






4 

1 1 1 I 1 



.10 r 



.08 - 



.06 - 



I- 

Q 

u 



.04 



.02 



(b) 



I 8a. n o 




-l_ 



-4 



12 16 



a,|,(deg) 



2 .4 .6 .8 1.0 



Figure 13. Variation of (a) lift and (b) drag coefficients for the RClO/05 tip. Re = 2.76 x 10^. 



42 



1.0 



.6 



a Ao=-io.o' 

O AO = -8.0'' 
A A0 = -2.5° 




(a) 



-4 



12 16 



a,,,(deg| 




Figure 14. Variation of (a) lift and (b) drag coefficients for the RClO/05 tip, Re = 3.91 x lO^. 



.05 



-.05 



e" -.10 



-.15 



-.20 



D Ae=-10.0° 
O AS =-8.0° 
A Afl=-2.5° 



-.25 



(a) 



\. 



N 




(b) 




-.2 



.2 



.8 1.0 



-.2 



.6 



.8 1.0 



Figure 15. Variation of pitching moment coefficient with lift coefficient for the RC 10/05 tip; 
(a) Re = 2.76 x 10^, (b) Re = 3.91 x 10^. 



43 



.05 



- 



-.05 - 



E -.10 



-.15 



-.20 



-.25 



D Afi=-10.0° 
O A5=-8.0° 
A Ae=-2.5° 




> \ 



\ 



\. 




(b) 



12 



16 



a,|,(deg) 



arp(deg) 



12 



16 



Figure 16. Variation of pitching moment coefficient with angle of attack for the RC 10/05 tip; 
(a) Re = 2.76 x 10^, (b) Re = 3.91 x 10^. 



44 



ORIGINAL" PAG^ 
BLACK mo WHITE PHOTOGRAPH 





Figure 17. Oil surface flow on the RClO/08 tip; ay = -5.0°, AG = -13.0°, Re = 3.91 x 10^; (a) upper 
surface, (b) lower surface. 



45 



ORIGINAL PAGE 
8UACK ^D WHITE PHOTOGRAPH 





Figure 18. Oil surface flow on the RClO/08 tip; ttj = -30°, AG = -13.0°, Re = 3.91 x 10^; (a) upper 
surface, (b) lower surface. 



46 



ORIGINAL PAGE 
BLACK mo WHITE PHOTOGRAPH 





Figure 19. Oil surface flow on the RClO/08 tip; aj = -1.0°, AG = -13.0°, Re = 3.91 x 10^; (a) upper 
surface, (b) lower surface. 



47 



ORIGINAL PAGE 
BLACK mo WHITE PHOTOGRAPH 





Figure 20. Oil surface flow on the RClO/08 tip; aj = 2.0°, AG 
surface, (b) lower surface. 



= -13.0°, Re = 3.91 x 10^; (a) upper 



48 



ORiGINAL" PAGE ' 
BLACK fmO WHITE PHOTOGSf\Ph 





Figure 21. Oil surface flow on the RClO/05 tip; ay = -2.5°, Ae = -2.5°, Re = 3.91 x 10^; (a) upper 
surface, (b) lower surface. 



49 



ORIGINAL PAGE 
BLACK mo WHiTE PHOTOGRAPH 





Figure 22. Oil surface flow on the RClO/05 tip; aj = 2.5°, AG = -2.5°, Re = 3.91 x 105; (a) upper 
surface, (b) lower surface. 



50 



ORIGINAL PAOE 
BLACK ^D WHITE PHOTOGRAPH 





Figure 23. Oil surface flow on the RClO/05 tip; aj = 7.5°, A0 = -2.5°, Re = 3.91 x 10^; (a) upper 
surface, (b) lower surface. 



51 



ORIGINAL PAGE 
BLACK mo WHITE PHOTOGRAPH 





Figure 24. Oil surface flow on the RClO/05 tip; aj = 12.5°, AG = -2.5°, Re = 3.91 x 105; (a) upper 
surface, (b) lower surface. 



52 



.75 r 



.50 - 



.25 



-.25 



Re=2.76X10^ 
Re=3.91X105 
Re=4.79X10^ 
Re=5.53X10^ 
Re=6.18X105 



(a) 



(b) 



6 8 

-Ae(deg) 



10 



12 



14 



6 8 

-AO(deg) 



10 



12 



14 



Figure 25. Tip aerodynamic center locations for (a) the RClO/08 tip and (b) the RClO/05 tip in a 
wing-fixed coordinate system, as a function of A9. 



nRe=2.76X10^ 
O Re=3.91X10 




-.05 



o 
E 



-.10 



4 6 8 10 

-AO(deg) 



14 



.15 




(b) 



4 6 8 

-AO(deg) 



10 



12 14 



Figure 26. (a) Lift and (b) pitching moment coefficients for the RC 10/08 tip as functions of AG; 
ttT = 0°. 



53 



.8 



o 



.4 



D Re=2.76X10^ 
O Re=3.91X10^ 
A Re=4.79X10^ 
+ Re=5.53X10^ 
X Re=6.18X10^ 



(a) 



-2 



-.05 




o 
E 



.10 



4 6 8 

-Ae(deg) 



10 



12 



14 



-.15 




(b) 



-2 



6 8 

-AO{deg) 



10 



12 



14 



Figure 27. (a) Lift and (b) pitching moment coefficients for the RC 10/05 tip as functions of A9; 
aj = 0°. 



54 



REPORT DOCUMENTATION PAGE 



Form Approved 
OMBNo. 0704-0188 



Public repc.ing burden tor ,his collecion o. information Is .s.ima.ed.o aver age 1 hooper f^^P""^'. '"^^^^''j^^^^^^ '£;.'^7^i 

garnering and maintaining the data needed, and complet ng and ™™«J7„*« =°'«?'°"^^^^^^^ and Reports, 1 215 Jefferson 



1. AGENCY USE ONLY (Leav» b/an*; 



2. REPORT DATE 



September 1991 



3. REPORT TYPE AND DATES COVERED 

Technical Memorandum 



4. TITLE AND SUBTITLE 



Experimental Study of an Independendy Deflected Wingtip 
Mounted on a Semispan Wing 



6. AUTHOR(S) 

D. M. Martin (University of Kansas Center for Research, Inc., 
Lawrence, Kansas) and L. A. Young 



7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES) 

Ames Research Center 
Moffett Field, CA 94035-1000 



9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES) 

National Aeronautics and Space Administration 
Washington, DC 20546-0001 



5. FUNDING NUMBERS 



RTOP 505-61-51 



8. PERFORMING ORGANIZATION 
REPORT NUMBER 



A-90210 



10 SPONSORING/MONITORING 
AGENCY REPORT NUMBER 



NASA TM- 102842 



11. SUPPLEMENTARY NOTES 

Point of Contact: Daniel M. Martin, Ames Research Center, MS T-042, Moffett Field, CA 94035-1000 
(415) 604-4566 or FTS 464-4566 



12a. DISTRIBUTION/AVAILABILITY STATEMENT 

Unclassified-Unlimited 
Subject Category - 02 



12b. DISTRIBUTION CODE 



13. ABSTRACT (Maximum 200 words) 

The results of a subsonic wind tunnel test of a semispan wing with an independently deflected tip 
surface are presented and analyzed. The tip surface was deflected about the quarter-chord of the 
rectangular wing and accounted for 17% of the wing semispan. The test was conducted to measure the 
loads on the tip surface and to investigate the nature of aerodynamic interference effects between the 
wing and the deflected tip. Results are presented for two swept tip surfaces of similar planform but 
different airfoil distributions. The report contains plots of tip lift, drag, and pitching moment for vanous 
Reynolds numbers and tip deflection angles with respect to the inboaid wing. Oil flow visualization 
photographs for a typical Reynolds number are also included. Important aerodynamic parameters such 
as lift and pitching moment slopes and tip aerodynamic center location are tabulated. A discussion is 
presented of the relationship between tip experimental data acquired in a steady flow and the prediction 
of unsteady tip motion at fixed-wing angles of attack. 



14. SUBJECT TERMS 

Free-tip rotor. Deflected tip, Experimental data 



17. SECURITY CLASSIFICATION 
OF REPORT 

Unclassified 



18. SECURITY CLASSIFICATION 
OF THIS PAGE 

Unclassified 



19. SECURITY CLASSIFICATION 
OF ABSTRACT 



15. NUMBER OF PAGES 

64 



16. PRICE CODE 

A04 



20. LIMITATION OF ABSTRACT 



NSN 7540-01-280-5500 



Standard Form 298 (Rev. 2-89) 

Prescribad by ANSI Std- Z39-1 8 
236-102