Vortex Flap Technology: a Stability and Control Assessment

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section, in the freestream direction madepossible the visualization of the

streamwise developmentof the vortex flow. Video equipment was used to I

record the flow development from a side view.

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The wind tunnel data acquisition system is portable and self-contained.

The data system uses a Cromemco Z-2, S-100 based computer system. Front-end

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signal conditioning is provided for all standard-type measuring devices such

as strain gages, pressure transducers, and thermocouples. The front-end

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hardware is computer-controlled to permit real time as well as post-test

data reduction. Both raw and reduced data are stored on floppy discs. The

computer software was already available for this test and other software is

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available for special applications. The computer supports a FORTRAN IV

Compiler for program development.

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Whenthe wind tunnel test was completed, the reduced data were trans-

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ferred from floppy discs to a 9-track magnetic tape. The 9-track tape was

then loaded into the IBMMain Computing System. The data were manipulated

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on the IMB system and plotted on a Tektronix 4014-1 terminal with hard copy

unit using existing stability and control derivative routines and plotting

packagesmodified for use in the NASA contract. I

The lateral-directional stability derivatives were obtained by subtract-

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ing the two spline-fitted rolling momentcurves (at zero and five degrees of

sideslip) at interpolated points, dividing by the constant sideslip angle, and

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spline-fitting the resulting curve. Similarly, the longitudinal stability

derivatives were obtained by determining Cmaand CNafrom their respective

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spline-fitted curves, dividing Cmaby CNaat interpolated values, and

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spline-fitting the resulting curve.

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DISCUSSION OF RESULTS

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The results of the wind tunnel test, including force and moment data, pressure measurements, and flow visualization, will be presented and dis-

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cussed in sections organized as follows:

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1) Effects of Deflected Vortex Flaps

2) Comparisons of Flap Effects

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a) Comparison of Full- and Part-Span Vortex Flaps b) Comparison of Tabbed and Plain Vortex Flaps

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c) Comparison of Conventional and Vortex Flaps

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3) Configuration Effects a) Trailing-Edge Sweep Variation Effects

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c) Nose Strake Effects d) Canard Effects

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e) Flap Apex Modification Effects

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Flap Deflection Angle Effects a) Leading-Edge Flap Effects

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b) Trailing-Edge Flap Effects c) Differential Flap Deflections for Roll Control

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Vertical Tail Effects

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a) Outboard Fin Effects

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b) Tail Deflection Effects

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Effects of Deflected Vortex Flaps

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The effects of a 30-degree deflection of the VLM-designed vortex flaps on the static longitudinal aerodynamic characteristics of the eleven

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basic wings examined are shown in Figures 14 through 24. In general, the designed full-span vortex flap results in a reduction in the lift

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coefficient at a given angle of attack, an increase in maximum lift, a decrease in drag-due-to-lift, and a nose-down increment in the pitching

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moment with little effect on the longitudinal stability level. Vortex flap deflection reduces vortex strength which in turn reduces the lift

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coefficient at a constant angle of attack. The increased CLmax trend is due to improved flow separation and vortex stability characteristics. The drag improvements are quite significant on the wings of moderate sweep; however,

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as the leading-edge sweep increases, vortex flap effectiveness decreases.

For example, at a typical maneuver CL of 0.5, the CD due to vortex flap

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deflection on the 45-degree delta wing is 45% whereas the CD for the same conditions on the 70-degree delta wing is 16%.

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The designed part-span vortex flap results in a decrease in lift

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coefficient below CLmax and slight nose down pitching moment increments at low-to-moderate angles of attack. The effect of part-span vortex flap

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deflection on induced-drag is minimal, due to the development of a multiple vortex system and adverse interaction between the respective vortex flows

which were observed in surface flow patterns and will be discussed in a I

later section.

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A plausible explanation for the reduced vortex flap effectiveness on the higher swept wings is as follows. The VLM design method yielded large

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flaps on the more highly-swept wings. Deflection of these large leading- edge control surfaces requires a much higher angle of attack to achieve a

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given lift. In addition, as wing sweep is increased the vortex is displaced away from the wing surface and quickly migrates off the deflected flap surface as angle of attack is increased. On the moderately-swept wings, the

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leading-edge vortex is better maintained on the flap surface to higher angles of attack than on the more highly-swept wings. This effect is

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apparent from wing upper surface flow patterns. Figures 25-27 are photo- graphs of the flow patterns on the 50-, 60-, and 70-degree cropped delta

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wings, respectively, taken at an angle of attack of 24 ° and zero sideslip.

It is evident that the vortex-induced primary reattachment lines move

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inboard as the leading-edge sweep is increased at a constant angle of attack.

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In addition, even under conditions where the vortex acts principally on

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the deflected flap, the aerodynamic thrust component due to vortex-induced suction pressures is smaller on wings of greater sweep. This is due to the more aft inclinations of the suction force which acts normal to the flap

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surface. Also, for the design flap condition of _n = 30°' the flap deflec- tion angle measured streamwise decreases with increasing wing sweep.

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The static longitudinal stability derivative, aCm/aCN, is plotted

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against angle of attack for the designed vortex flapped configurations in Figures 28-38. Designed Vortex flaps delay to higher angles of attack the

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stable break in the longitudinal stability curve. However, designed flaps generally do not affect the character of the stability curve.

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Coupling between the longitudinal and lateral-directional axes is

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represented by the parameter Cm]#[ , which is indicative of the tendency to pitch due to sideslip. It is desireable to have as small a positive value of this parameter as possible or a negative value to avoid increasing the

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angle of attack by pitching up in sideslip. A negative slope of the Cml#l curve in the stall or post-stall region is desireable to promote stall

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recovery.

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The variation of the pitch-sideslip coupling parameter with angle of attack is shown in Figures 39-49. Vortex flap deflection reduces the

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nose-up pitching moment due to sideslip and reduces the slope of the curve at angles of attack greater than 10° . These effects are less pronounced on

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development from the undeflected portion of the wing leading-edge.

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The effects of vortex flaps on the static lateral-directional stability

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derivatives, ac_/a_, aCN/a #, and aCy/a#, are shown in Figures 50-60.

With undeflected flaps the wings exhibit large variations in rolling moment

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at high angles of attack and consequent fluctuations in lateral stability,

produced by asymmetric vortex breakdown in sideslip. The unstable break

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in the lateral stability curve occurs at higher angles of attack with

increasing wing sweep. At higher attitudes, the stable variation of

C_n with c can be attributed to more symmetric wing flow separation I

w characteristics. In this region, the wings of moderate sweep exhibit massive, unsteady flow separation and, therefore, fluctuations in the

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rolling moment. This unsteady flow situation was especially noticeable during wing upper surface flow visualization. The kerosene was very slow

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in evaporating, indicative of low-energy flow. The surface flow patterns were ill-defined and exhibited abrupt fluctuations at fixed angle of attack.

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In contrast, the surface patterns on the more highly-swept wings were characterized by a strong rotational flow, although vortex bursting occurred well upstream of the wing trailing-edge.

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At low angles of attack, designed vortex flaps reduce lateral stability

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due to vortex separation on the flap underside and provides stable increments to the lateral stability at moderate angles of attack as a result of reduced

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tip separation. At high angles of attack the designed vortex flaps delay the onset of vortex breakdown and reduce burst asymmetries. These effects

serve to delay the unstable break in and diminish the large oscillations of I

the C_# curve.

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These trends were not exhibited by the more highly-swept 70-degree cropped delta and 76.5/66.5-degree cranked wings. Vortex flap deflection

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has a destabilizing effect throughout the angle of attack range. The vortex flaps on these configurations represent a large percentage of the total wing

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area. Deflection of these large surfaces suppresses vortex development at low angles and significantly reduces vortex strength at higher angles of

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attack, which reduces the magnitude of C_.

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At low angles of attack, leading-edge vortex flap deflection contributes small stable increments to the directional stability due to vortex suction

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pressures acting on the back-side of the flap. At moderate angles of attack the deflected flap is destabilizing directionally on the more highly-swept

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configurations. This may be the result of reduced vortex-tail interaction since the wing vortical flow is of reduced strength and is displaced

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outboard with flap deflection. Vertical tail-off tests would be required to confirm this hypothesis. At high angles of attack, the unfavorable sidewash

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which results from vortex bursting over the wing adversely affects the directional stability although the fluctuations in the unstable yawing moments are reduced by the deflection of the vortex flaps.

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Vortex flap deflection induces small variations in the sideforce due to

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sideslip at the low-to-moderate angles of attack for all but the most moderately swept. At the high angles of attack the sharp changes in the

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slope of the sideforce curves are reduced by vortex flap deflection on the moderately swept wings. On the most highly-swept wings, vortex flap deflection

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induces a positive increment in Cy# at all angles of attack, consistent with reduced load on the vertical tail.

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Figures 61-63 present the variations of lift and rolling moment with sideslip at several angles of attack on the 50-degree and 65-degree cropped

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delta and 70/50-degree cranked wings. The 50-degree cropped delta wing exhibits a nominal variation of lift with sideslip up to 24 degrees as shown

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in Figure 61. The rolling moment variation with sideslip is stable at all angles of attack, although the magnitude of the rolling moment coefficient

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decreases with increased angle of attack. The moderately swept wing does not generate a strong vortex system and is therefore not sensitive to vortex

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burst asymmetries and the consequent aerodynamic nonlinearities at high angles of attack.

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Figures 62 and 63 show similar variations of lift and rolling moment

with sideslip on the 65-degree cropped delta and 70/50-degree cranked

wings. At the higher attitudes, the stable rolling momentcoefficients are

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greater in magnitude in comparison to the 50-degree swept wing due to the

increased vortex-induced effects. Sideslip "sweeps" were not performed at

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sufficiently high e's to incur pronounced nonlinearities associated with

asymmetric core breakdown.

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: 50-Degree and 65-De_ree Dynamic Directional Stability Parameter, C n_dy n Cropped and 7(I/50-De_ree Cranked Wings

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The parameter C n (see Reference 11) is a measure of the direc-

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#dyn tional stability about the flight path under dynamic conditions. It is

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computed from: m sin (%.

Cn Cn I

= cosa - Iz C£# #dyn # Ix

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Generally used to predict yaw departure, negative values of Cn#dy n

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indicate the possibility of directional divergence. A moment-of-inertia ratio (Iz/I x) of 8 was used as a representative value for an advanced

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tactical fighter configuration in calculating the dynamic directional stability parameter.

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I Figures 64-66 present the effect of a 30-degree deflection of the

vortex flaps on Cn for the 50-degree and 65-degree cropped delta and _dyn

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70/50-degree cranked wings. The 50-degree cropped delta wing maintains positive values of Cn up to the maximum angle of attack tested.

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_dyn Vortex flap deflection increases the positive values of the dynamic

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directional stability parameter in the angle of attack range of 7 to 22 degree but reduces Cn significantly at higher angles of attack.

_dyn

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Vortex flap deflection has no effect on the angle of attack at which the minimum value of Cn occurs.

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_dyn

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Figure 65 shows positive values of Cn for the 65-degree cropped _dyn

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delta wing up to an angle of attack of 27 degrees with vortex flaps unde- flected and to 29 degrees with flaps deflected. Vortex flap deflection

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results in pos!tive increments to Cn at angles of attack greater _dyn than 20 degrees. However, the large negative values of the parameter at

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angles of attack greater than approximately 29 degrees indicate the config- uration is susceptible to yaw departure.

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The configuration with the 70/50-degree cranked wing exhibits large,

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negative Cn values above 29 degrees angle of attack, as shown in #dyn Figure 66. This effect is mitigated somewhat by vortex flap deflection.

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Similar to the results obtained on the 65-degree cropped delta wing, the static lateral stability derivative (Figure 58a) is the primary driver of

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the Cn parameter.

#dyn

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upper Surface Static Pressure Distributions: 65 Degree Cropped Delta Win 9

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Figures 67 and 68 present the effect of sideslip angle on the 65-degree cropped delta wing upper surface static pressure distributions at a= 16

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degrees and e= 24 degrees. Vortex flap deflection angle is fixed at 30 degrees. Upper surface static pressure coefficient, CPU, is plotted against

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span distance, Y (PORT), normalized by the vortex flap hingeline span distance, S (FLAP HINGELINE). Pressures outboard of a non-dimensional span

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location of 1.0 are on the deflected vortex flap; pressures inboard of this position are on the main wing.

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Results at a=16 degrees generally reveal a reduction in the peak vortex-induced suction pressures on the windward and leeward wings as

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sideslip angle is increased. The location of the peak pressures on the windward wing is sensitive to sideslip angle, displaying an inboard shift

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due to sideslip. However, the peak pressure magnitudes on the windward and leeward wings are comparable at a given measurement station.

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Flow-field as_nnmetries due to sideslip are more pronounced at e=24

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degrees. A consistent reduction in the peak suction pressures occurs on the windward wing. In fact, the absence of a pronounced peak in the pressure distributions at the aft measurement station is indicative of vortex break-

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down. In contrast, the peak vortex-induced suction pressures on the leeward wing increase at the forward measurement station and are insensitive

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to sideslip at the mid- and aft stations. In addition, there is a marked outboard shift in the location of the peak pressures. These flow-field

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characteristics correlate well with the reduced lateral stability and with the variation of lift and rolling moment with sideslip shown previously in

Figures 54 and 62, respectively, at similar angles of attack. I

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upper Surface Static Pressure r)istributions: 70/50-Degree Cranked Wing

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Figures 69 and 70 illustrate the effect of sideslip on the cranked wing

I surface pressures ate=16 degrees and a=24 degrees, respectively. Upper

surface static pressure distributions on the windward and leeward wings are presented at two measurement stations ahead of the leading-edge break

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(x/c=.422 and x/c=.678) and one station aft of the break (x/c=.848). Vortex flap deflection is fixed at 30 degrees.

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Sideslip has only a small effect on the wing surface pressures at a=16

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degrees. On the windward wing, the magnitude of the peak suction pressures induced by the vortical flow shed from the more highly-swept inboard leading edge is not sensitive to sideslip angle. However, the locations of the peak

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pressures ahead of and behind the planform break are more outboard and inboard, respectively, at sideslip. The reduced suction peak on the deflec-

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ted flap of the outboard wing panel is the result of an effective increase in the streamwise deflection and associated reduction in vortex strength.

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On the leeward wing, the peak suction pressures are generally lower in magnitude as the sideslip angle is increased. This is indicative of a

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displacement of the leading-edge vortex away from the wing surface.

Similar effects of sideslip on the wing surface pressures are observed

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at a=24 degrees. The test results indicate that the maximum suction pres- sures are maintained on the deflected flap at the forward measurement

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station on the windward and leeward wings. Migration of the vortex off the flap is evident at the mid station on both wings. In addition, a small,

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concentrated vortex is maintained on the outboard vortex flap segments. The pressure distributions provide no indication of vortex breakdown effects.

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In fact, the surface pressure data trends correlate well with the high level of lateral stability exhibited by the cranked wing at a=24 degrees.

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Comparison of Flap Effects

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Comparison of Full- and Part-Span Vortex Flaps

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Full- and part-span vortex flaps were designed for the 65- and 70- degree cropped delta and the 70/50-degree cranked wings, as shown in Figures

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71-73. Comparisons of the static longitudinal aerodynamic and stability characteristics of the three wing planforms with full- and outboard part-

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span flaps deflected 30 degrees are shown in Figures 74-79.

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The deflected full-span flap results in higher maximum lift, lower induced drag, and less nose-up pitching moments than the part-span flap on the 65-degree cropped delta wing. These trends also apply to the 70-degree

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cropped delta wing, although it was not tested up to CLmax.

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In contrast, the 70/50-degree cranked wing with part-span flap exhibits higher lift throughout the range of angles tested and lower drag-due-to-lift

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at high lift coefficients, relative to the full-span flap. These results are consistent with wing surface pressure measurements which revealed the

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development of a strong vortex from the inboard, undeflected portion of the wing leading-edge. This vortex induced an effect analogous to a wing leading-edge extension (LEX). Similar to the cropped delta wing results,

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the part-span flap promotes nose-up pitching moments in comparison to the configuration with full-span flap.

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The effects of flap deflection on the coupling parameter, Cm/#/,

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plotted against angle of attack, are show in Figures 80-82. The full-span vortex flaps result in less nose-up pitching moment due to sideslip for

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all configurations.

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Comparisons of the effects of 30 ° full- and part-span vortex flap

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deflections on the lateral-directional stability characteristics of the three configurations are shown in Figures 83-85. The results are somewhat configuration dependent. Compared to the part-span flap, the full-span flap

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on the 65-degree cropped delta wing results in significantly higher lateral stability above 10 degrees angle of attack and delays to higher angles the

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unstable break in the stability curve. In contrast, on the 70-degree cropped delta and the 70/50-degree cranked wings the full-span flaps exhibits

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reduced lateral stability relative to the part-span flap throughout most of the angle-of-attack range. However, the full-span flap on the 70-degree

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cropped delta wing eliminates the large oscillations in lateral stability exhibited by the wing with part-span flap.

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The comparisons show no significant differences between the effects of the full- and part-span flaps on the directional stability characteristics

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nor on the sideforce due to sideslip. In general, the wings with part-span flaps are slightly more stable in yaw at low angles of attack, whereas

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the full-span flap tends to delay the onset of directional instability to slightly higher angles of attack.

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Figures 86, 87 and 88 are photographs of upper surface flow patterns on the 65-degree and 70-degree cropped delta and 70/50-degree cranked wings

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with full- and part-span flaps, respectively. All patterns correspond to a flap deflection of 30 degrees at an angle of attack of 12 degrees and zero

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sideslip. The photos show that a dual-vortex system is generated by the part-span flap configuration: one vortex develops from the undeflected

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inboard portion and another emanates from the deflected flap.

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The local increase in vortex strength along the undeflected leading- edge accounts for the increased non-linear lift in the moderate angle of attack range and the nose-up pitching moment increments associated with the

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part-span flaps. The single concentrated vortex emanating from the deflect- ed full-span flap induces a greater thrust component since the vortex acts

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on a larger forward facing surface area, which accounts for the lower induced-drag associated with the full-span flap.

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Due to the shorter generating length of both the flap vortex and the

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inboard vortex, the part-span flap configuration develops more pronounced

vortex bursting at the moderate-to-high angles of attack. This is also an

explanation for the reduced maximum lift typically exhibited by the part-

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span flap configurations. In addition, the wings with part-span flaps are

more prone to vortex-burst asymmetries in sideslip at moderate angles of

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attack.

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It should be noted that the lateral stability characteristics of the

wings with part-span flaps are sensitive to small changes in the flap apex

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geometry. Consequently, the trends established in Figures 74-85 are not

necessarily universal. This topic will be addressed in more detail in a

later section.

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Comparison of Tabbed and Plain Vortex Flaps

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Comparisons of the longitudinal aerodynamic and stability characteris- tics of the 6(}- and 65-degree cropped delta and the 70/50-degree cranked

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wings with tabbed and plain vortex flaps deflected 30 degrees, illustrated in Figures 89-91, are shown in Figures q2-97. In general, the wings with tabbed vortex flaps exhibit higher lift and lower drag at high lift. These

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results are indicative of the increased vortex strength arising from the controller tabs. The tabs had no effect on the longitudinal stability

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characteristics of the cropped delta wings and only slightly increased the pitch instability of the cranked wing at high lift.

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The comparisons in Figures 98-100 indicate that pitching moment due to

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sideslip is insensitive to this variation in flap geometry.

Figures 101-103 present comparisons of the lateral-directional sta-

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bility characteristics. The plain vortex flap on the 60-degree cropped delta wing results in significantly higher levels of lateral stability in

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comparison to the wing with the tabbed vortex flap throughout most of the

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angle of attack range. This trend also applies to the 65-degree cropped delta and 70/50-degree cranked wings, although the effect is considerably

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maintain lateral stability at high angles of attack while the plain vortex flapped-configurations become unstable. Although data were not obtained on

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the tabbed flaps for comparison to the plain flaps at the highest angles it appears that this trend would be maintained.

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There is little difference in the respective directional stability

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characteristics although the tabbed flap is somewhat less unstable than the plain vortex flap at high angles of attack.

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The laser light sheet technique was applied to the 65-degree cropped delta wing with both flap configurations to analyze the respective flow

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field structures. Results from the flow visualization studies indicate that the tabbed vortex flap configuration was more effective in trapping the

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vortex along the wing leading-edge than was the plain vortex flap at low angles of attack. However, the vortex was maintained on the plain vortex

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flap to higher angles of attack, particularly in the .4 to .6 CL range, than on the tabbed vortex flap.

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Upper Surface Static Pressure Distributions With Tabbed Vortex Flaps:

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65-Degree Cropped Delta and 70/50-Degree Cranked Wings

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Figures 104 and 105 show the effect of sideslip angle on the 65-degree cropped delta and 70/50-degree cranked wing pressure distributions with

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tabbed vortex flaps deflected 30 degrees. The cropped delta wing results obtained ata=12 degrees reveal nominal effects due to sideslip on the magnitude and location of the peak suction pressures on the windward and

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leeward wings. The vortex is effectively "captured" at this angle of attack by the unique tabbed flap geometry. Similar results are obtained on the

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cranked wing, although the maximum vortex-induced suction levels generally decrease with sideslip.

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Comparison of LIpper Surface Static Pressure Distributions with VLM-Designed

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and Tabbed Vortex Flaps: 65-Degree Croped Delta and 70/50-Degree Cranked Wings

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Figures 106 and 107 compare the pressure distributions on the 65-degree cropped delta wing with VLM-designed and tabbed vortex flaps at a=8 degrees

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and a=12 degrees, respectively. A similar comparison at a=12 degrees is presented in Figure 1N8 corresponding to the 70/50-degree cranked wing.

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The results obtained on the cropped delta wing reveal consistently

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higher suction levels at the forward and middle pressure measurement sta- tions with the tabbed vortex flap. The data trends show the effectiveness

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of the controller tabs to manipulate the leading-edge vortex strength at a given angle of attack.

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The test data comparisons on the cranked wing reveal more pronounced variations in the pressure distributions. The stronger vortex on the wing

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with tabbed flap is evidenced by the significant increase in the suction levels on the deflected flap at all measurement stations. In addition, the

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larger (and stronger) vortex on the outboard tabbed flap segment induces primary flow reattachment near the hingeline, thereby alleviating flow

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separation in this region.

The pressure data trends correlate well with increased lift obtained in

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force measurements and more concentrated vortex observed in laser lightsheet flow visualization of the tabbed-flap configurations.

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Vortex Flap Planform Effects

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Vortex flaps of conventional taper were empirically designed for the 60-, 65-, and 70-degree cropped delta and the 70/50-degree cranked wings for comparison to VLM-designed vortex flaps of equal exposed area. The eight

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configurations are shown in Figures 109-112. The term "conventional" is

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somewhat misleading since it implies attached flow as opposed to vortex flow. It is apparent from surface flow visualization that flaps of conven-

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dominated at moderate-to-high angles of attack. However, because of the difference in flap chord variation, the flow situations are quite

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different. The VLM-design procedure provides a flap shape, typically of inverse taper, corresponding to the flow condition of vortex separation

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everywhere along the leading-edge with primary reattachment at the hingeline of the deflected flap. The tapered planform, on the other hand, tends to

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have regions of vortex-dominated flow with reattachment non-coincident with the hingeline. Flaps of sufficient taper will also feature mixed regions of attached flow inboard and vortex flow outboard.

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Comparisons of the effects of VLM-designed and tapered flaps deflected

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30 degrees on the longitudinal aerodynamic and stability characteristics of the four wing planforms are shown in Figures 113-120. The VLM-designed

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vortex flaps yield consistently lower lift at a given angle of attack and higher drag in comparison to the tapered flaps. The wings with VLM-designed

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vortex flaps also exhibit consistently greater pitch instability and more nose-up pitching moment due to sideslip.

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Figures 121-124 present comparisons of the lateral-directional sta- bility characteristics. The VLM-designed flaps deflected to 30 degrees

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generally result in a higher level of lateral stability than the con- ventionally-tapered flaps. This effect is due to the larger flap chord

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outboard which delays tip stall to higher angles of attack. This trend does not apply to the 70-degree cropped delta wing, however. Wing surface

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pressure measurements obtained on the wing with VLM-designed vortex flap indicate that the pronounced inverse taper flap planform significantly reduces vortex strength near the tip. The results from the 70-degree swept

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wing suggest that lateral stability favors a stronger vortex outboard, as is the case with the tapered flap.

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For all the configurations examined, the tapered flaps result in lower

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directional stability at low-to-moderate angles of attack and greater instability at high angles of attack.

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Configuration Effects

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Trailing-Edge Sweep Effects

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Vortex flaps were designed for the 65-degree swept wing with trailing-

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edge sweep angles of 0 °, +_15 ° , and -30 ° , as shown in Figure 125. Neither the 15-degree forward sweep nor the 15-degree aft sweep had a significant

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effect on the size of the vortex flap required to meet the VLM-design criteria. However, 30-degree forward sweep required an 8-10% increase in the flap local chord length. All flaps were designed for a 30-degree

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deflection angle.

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The static longitudinal aerodynamic and stability characteristics for the range of trailing-edge sweep are shown in Figures 126-127. Aspect

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ratio appears to be the dominant effect on lift, with the cropped arrow wing developing higher lift throughout the angle of attack range tested.

The markedly improved drag polar associated with the cropped arrow

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(A.TE = +15 ° ) wing is the result of higher aspect ratio and the larger ratio of vortex flap-to-wing area. The cropped arrow wing exhibits reduced

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pitch instability. This can also be attributed to the effectively larger vortex flap and to a slight forward shift in the .40c location.

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Consistent with these trends, the cropped arrow planform with vortex

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flap deflected 30 degrees develops less nose-up pitching moment due to sideslip. This result is shown in Figure 127.

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The effects of trailing-edge sweep variation on the lateral-directional

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stability characteristics are shown in Figure 128. The cropped diamond planforms exhibit reduced lateral stability at low-to-moderate angles of

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attack relative to the baseline cropped delta configuration. In contrast, the cropped arrow configuration generally improves lateral stability in the

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same angle of attack range. The latter, however, shows considerably reduced levels of lateral stability at higher angles of attack due to the more severe vortex breakdown asymmetry that is characteristic of arrow wings.

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The higher stability levels associated with the cropped arrow wing are probably due to an increase in tail volume, which results from the forward shift of the reference c.g. and reduced wing area, rather than an aerodyna-

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mic effect. Similarly, the reduced directional stability of the diamond wing is probably due to a reduction of tail volume because of the aft shift of the reference c.g. and an increase in wing area.

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Win 9 Vertical Position Effects

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The 65-degree cropped delta wing with full-span vortex flaps deflected 30 degrees was tested at three non-dimensional wing vertical positions,

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z/d = -0.20, 0.0, +0.20, with z measured normal to the model centerline and non-dimensionalized by the maximum body width, d.

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The longitudinal aerodynamic and stability characteristics with the high-, mid-, and low-wing positions are shown in Figures 129-130. Relative

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to the mid-wing position, the high- and low-wing positions result in in- creased lift, less induced-drag at mid-to-high lift, and slight nose-up

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pitching moment increments with no change in longitudinal stability.

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Wing upper surface flow visualization, however, revealed no discernible changes in the flow patterns due to wing height variation. Specifically, the vortex-induced primary reattachment and secondary separation lines

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appeared insensitive to wing position on the body.

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Figure 131 illustrates the effect of wing vertical position on the

I

lateral-directional stability characteristics. Comparedto the baseline

mid-wing position, the low-wing and high-wing configurations exhibit lower

I

and higher levels of lateral stability, respectively. These effects on

the lateral stability are due to the different flow fields the configur-

I

ations experience due to the crossflow around the fuselage. In sideslip,

the low wing configuration experiences downwash on the windward wing and

upwashon the leeward wing, while the high-wing configuration experiences

l

upwashon the windward wing and downwash on the leeward wing. These situ-

ations contribute positive and negative rolling momentincrements to the

!

baseline conditions, respectively, consistent with the classical fuselage

crossflow effect on lateral stability. The unstable break in the lateral

I

stability curve occurs at virtually the sameangle of attack for all three

configurations.

I

At low-to-moderate angles of attack there is no significant differ-

ence in the directional stability of the low-, mid-, and high-wing config-

!

urations. At moderate-to-high angles of attack, the high-wing results in

lower yaw instability while the low-wing position results in increased

I

instability. The variations in directional stability due to wing position

can once again be attributed to fuselage crossflow effects rather than any

!

changes in the vortex flap aerodynamics.

I

Effect of Wing Vertical Position on Upper Surface Static Pressure

Distributions: 65-Degree Cropped Delta Wing i

For completeness, pressure distributions are presented in Figure 132

I

corresponding to mid- and low-wing positions at a=16 degrees with vortex flap deflected to 30 degrees. The magnitude and location of the peak

I

vortex-induced suction pressures are relatively insensitive to wing vertical location on the fuselage. These results are consistent with upper surface

I

flow visualization which revealed no measurable changes in the vortex migratory behavior due to wing height variation.

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Nose Strake Effects

!

The nose strakes shown in Figure 133 were empirically designed and mounted at the nose maximum half-breadth. The strakes were fabricated from

I

.05-inch thick sheet metal and were joined by a thin strip of metal. The strip joining the strakes was fitted into a slot in the nose of the model

I

such that the strakes were flush with the surface with no protuberances.

The nose strakes were tested in conjunction with the 65-degree cropped delta

I

and 70/50-degree cranked wings. Vortex flap effects were examined on the cranked wing with deflection angles of 0 and 30 degrees; the cropped delta

I

wing was tested with a 30 degree flap deflection. Trailing-edge flaps were unde flected.

I

The effect of vortex flap deflection on the longitudinal aerodynamic and stability characteristics of the cranked wing with nose strakes is

I

presented in Figures 134 and 135. Based on a comparison with vortex flap effects on the same configuration without strakes, shown previously in Figures 22, 36, and 47, it is evident that the strakes do not alter the

I

vortex flap effects on the lift and drag characteristics. However, they do cause more nose-down pitching moment increments and a more significant

i

reduction of the pitch-sideslip coupling parameter due to flap deflection.

I

Figure 136 shows the effect of vortex flap deflection on the lateral- directional stability characteristics of the configuration with nose strakes

I

for comparison to the effects on the configuration without strakes, given in Figure 58. Both configurations yield similar trends in the lateral stability

I

although the strakes-off configuration results in more improvement in C_# at moderate-to-high angles of attack due to vortex flap deflection.

Although the strakes-on configuration generates higher levels of directional

I

stability than does the strakes-off configuration, vortex flap deflection has a more detrimental effect on the former.

I

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Figures 137-140 illustrate the effects of the strakes on the static

longitudinal aerodynamic and stability characteristics of the 65-degree

I

cropped delta and 70/50-degree cranked wings, both with vortex flaps de-

flected to 30 degrees. The strake-induced lift produces nose-up pitching

momentincrements, increased pitch instability, and more nose-up pitching I

momentdue to sideslip. However, the strakes have minimal effect on total

lift and drag.

I

Figures 141 and 142 showthe effects of the nose strakes on the

I

lateral-directional stability characteristics. The strakes result in

increased lateral stability at moderate-to-high angles of attack. In

addition, the strakes enable directional stability to be maintained to

I

higher angles of attack and considerably reduce the directional instability

at high angles of attack.

I

The effect on C n of nose strakes in conjunction with the 70/50-

#dyn

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degree cranked wing with vortex flaps deflected to 30 degrees is illustrated

in Figure 143. Consistent with the improved lateral-directional stability

I

shown previously in Figure 142a, the strakes result in positive values of

the dynamic directional stability parameter up to approximately 34 degrees.

I

To isolate the effect of the nose strake flow field, a component

build-up was conducted in conjunction with the 70/50-degree cranked wing. I

Figures 144 and 145 provide someinsight into the contributions to lateral-

directional stability of the various airframe components. The test data

I

suggest that a favorable interaction of the strake vortex system with the

windward wing flow field is the primary source of the increased dihedral

I

effect. At high angles of attack, a slight increase in windward wing vortex

stability can promote a large increase in lateral stability. This effect

I

has been observed in several investigations of advanced fighter aircraft

configurations. The main contribution to increased directional stability is

the direct suction effect induced by the windward strake vortex. Secondary

l

effects due to strake vortex system interaction with the centerline tail may

also contribute to the enhanced stability. Increased lateral and directional

I

stability due to nose strakes located at the maximum half-breadth has been

demonstrated in wind tunnel and flight tests of the Northrop YF-17

I

(Reference 12).

I

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I

Canard Effects

I

A pair of canards were empirically designed on the basis of results obtained in References 13 and 14, and were tested in conjunction with the

I

65-degree delta wing with full-span vortex flaps deflected 30 degrees.

Canard area was not included in the reference area. The objective of this

I

portion of the wind tunnel investigation was not to optimize canard-wing interactions but, rather, to obtain representative results of canard effects

I

on a vortex-fl apped wing.

I The canards, illustrated in Figure 146, were tested at a non-dimen-

sional height, h/b, of 0.085 above the wing plane, where h and b are the canard height and wing span, respectively. Two longitudinal positions, _,

I

non-dimensionalized by the wing mean aerodynamic chord, c, were investi- gated and corresponded to _/c=.40 and _/_=.27. Here, _ was defined

I

as the distance between the canard apex and the wing apex. The canard incidence angle was 0 degrees.

I

The effects of the canards on the static longitudinal aerodynamic and

I stability characteristics are presented in Figures 147-148. The canards

promote a more pronounced nonlinearity of the lift curve at high angles of attack due to canard-wing vortex interaction, induced-drag reductions at

i

high lift, large nose-up pitching moment increments, and increased pitch instability. Longitudinal trim conditions will require canard deflections

t

which may affect these results which are based on zero canard deflection.

I

Figure 149 presents canard effects on the lateral-directional stability characteristics. Canard downwash improves the windward wing flow separation

I characteristics and, consequently, results in large increases in lateral

stability at moderate-to-high angles of attack. The data suggest that the canards delay the onset of asymmetric vortex breakdown on the wing and

I

consequently delay the corresponding unstable break in the lateral stability curve. Concurrent with these favorable effects are unstable increments in

I

directional stability at low-to-moderate angles of attack. The canard alone

I

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I

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would be stabilizing directionally based on the sideforce due to sideslip

!

data which indicates that the canards induce positive sideforce increments.

I

Figure 150 showsthe effect of canards on the variations of lift and

rolling momentwith sideslip on the 65-degree cropped delta wing. The

canards have negligible effect on the lift loss at sideslip. However, the

I

data indicate that the canards result in a more stable variation of rolling

momentwith sideslip ata=24 degrees.

I

Figure 151 showsthe effect of the effect of canards on the dynamic

I

directional stability parameter for the 65-degree cropped delta wing. Above

20 degrees angle of attack the canards result in large positive increments

I

to C n . This effect is due largely to the improved lateral stability charac-

#dyn

teristics arising from a favorable canard-wing flow field interaction.

I

Figure 152 presents the effect of the canard on the wing upper surface

I

static pressure distributions at an angle of attack of 24 degrees. Only the

right wing was instrumented with pressure orifices. Consequently, test data

I

were obtained at positive and negative sideslip angles to access canard

effects on windward and leeward wing surface pressures, respectively.

i

The addition of the canards results in an outboard shift of the peak

vortex-induced suction pressures at all measurementstations, indicating an

improvement in the migratory behavior of the wing leading-edge vortex. For

example, with canards on, the vortex remains on the flap up to approximately

I

the mid-measur_ent station (x/c=.576). With canards off, the vortex has

migrated off the flap upstream of the first station (x/c=.405). The latter

I

position is inboard of the canard tip and, consequently, exhibits lower peak

suction pressures due to canard downwasheffects. The mid- and aft-pressure

measurementstations are outboard of the canard tip and generally develop

I

slightly higher suction levels as a result of the canard-induced upwash

fi el d.

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The canard has a pronounced effect on the wing upper surface pressure distributions in sideslip. Results are presented in Figure 153 correspond-

I

ing toa=24 degrees and #=10 degrees. The pressure distributions on the windward wing (Figures 153a -153c)exhibit a significant increase in the magnitude of the peak suction pressures and an outboard shift in the

i

position of the suction peaks at all measurement stations. These results are indicative of enhanced wing leading-edge vortex stability and downward

I

displacement of the vortex core due to the canard-induced downwash field.

In contrast, the canard generally decreases the vortex-induced effects on

I

the leeward wing (Figures 153d - 153f), presumably because of premature "lifting away" of the vortex from the wing surface and possible core break-

I

down due to canard upwash. The pressure distributions correlate well with the large increase in lateral stability shown previously in Figure 146.

I

Canard deflection necessary for longitudinal trim would certainly affect the interaction of the canard flow-field with the wing and tail

I

surfaces and therefore will affect the surface static pressure distributions as well as the lateral-directional stability trends. The results presented

I

here should be interpreted accordingly since zero canard incidence is unlikely, particularly at high angles of attack.

I

Vortex Flap Apex Modification Effects

I

Modifications were made to the apex of the part-span vortex flaps on

!

the 65-degree cropped delta and the 70/50-degree cranked wings. A stream- wise cut was made in the flap and the curved apex region inboard of the cut

I

was removed, as shown in Figure 154. Test results were obtained with vortex flaps deflected to 30 degrees. The data are not corrected for the slight

I

change in wing area.

The effects of these modifications on the static longitudinal aero-

I

dynamic characteristics are shown in Figures 155 and 156. In general, the wings with unmodified flaps exhibit higher lift, less drag at high lift, and

I

more nose-up pitching moments. The unmodified vortex flap on the 65-degree cropped delta wing results in higher maximum lift. This trend appears to

I

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be applicable to the 70/50-degree cranked wing, although the latter was not

I

tested up to C . Surface flow patterns indicate that the modified apex

Lmax

is conducive to unorganized flow separation just downstreamof the wing-flap I

junction and to the development of a counter-rotating vortex along the

streamwise edge. In contrast, the unmodified apex promotes the smooth

I

development of a leading-edge vortex from the deflected flap surface. These

effects, in combination with the reduced flap area and earlier vortex

I

bursting, account for the higher drag and lower lift associated with the

modified vortex flap at high a°s.

I

Figures 157 and 158 illustrate the effects of the flap modifications on the longitudinal stability characteristics. The flap apex modification has

I

a minimal effect on the longitudinal stability level of the cropped delta wing. However, the modified flap increases pitch stability on the cranked

I

wing. The cropped delta wing exhibits less nose-up pitching moment due to sideslip with the modified flap; nose down pitching moments due to sideslip

I

are developed on the cranked wing.

I

Figures 159 and 160 present the lateral stability characteristics of the 65-degree cropped delta and 70/50-degree cranked wings with the original and the modified part-span vortex flaps deflected to 30 degrees. The

I

modified flap on the cropped delta wing exhibits significantly higher levels of lateral stability at moderate and high angles of attack and a delayed

I

onset of the unstable break in the C_# curve. In contrast, the modified

I

flap on the cranked wing results in lower, although still favorable, lateral stability. However, the modified flap maintains lateral stability up to the

!

highest angles of attack tested while the original flap becomes unstable.

Surface flow patterns on the 65-degree cropped delta wing indicate that

I

the streamwise edge of the modified part-span flap appears to reduce the interaction of the flap vortex with the vortical flow that is generated by

I

the undeflected portion of the wing leading-edge. Wing upper surface static pressure distributions obtained in another investigation on the 65-degree

I

cropped delta wing with unmodfied flap indicate that at moderate-to-high

I

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I

angles of attack the vortices merge. This single, concentrated vortex

I

system is prone to large asymmetries in the core breakdown positions. The modified flap, however, enables a distinct two-vortex system to be maintained to higher angles of attack. Although the vortices break down sooner due to

l

their relatively short generating lengths, the vortex flow-field exhibits reduced breakdown asymmetries in sideslip. This effect is analogous to wind

I

tunnel and water tunnel results obtained on an advanced LEX-wing fighter configuration in Reference 15. In the latter study it was found that

!

truncating the LEX apex region decreased C but enhanced high-lateral Lmax

I

stability as a result of reduced vortex burst asymmetries.

An understanding of the conflicting results obtained on the cranked

I

wing is again aided by wing surface pressure measurements obtained in another study are shown in Figures 161 and 162. In contrast to the part-span

I

flap on the 65-degree cropped delta wing, the part-span flap on the cranked wing is not an efficient vortex generator. Although the sweep angle discon-

I

tinuity is not large, pressure measurements indicate that the variation of flap chord across the span is sufficient to promote a two-vortex system on

I

the deflected flap. However, only the vortex shed from the inner flap segment is of any consequence. The flap apex modification serves to reduce the generating length of this vortical flow. The beneficial effect on roll

I

stability of a strong vortex outboard is thereby reduced.

I

The original and modified flaps result in similar directional stability trends for both wing planforms, as shown in Figures 163 and 164.

I

Two apex modifications were made to the full-span vortex flap of the

|

60-degree cropped delta wing as shown in Figure 165. In one case, the inboard 25 percent of the exposed vortex flap span was removed and the remaining portion was deflected to 30 degrees. In the other case a stream-

I

wise cut was made in the flap at a distance of 25 percent of the exposed half-span. The inboard segment was undeflected while the outboard segment

I

was deflected to 30 degrees.

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Figures 166 and 167 illustrate the effects of the former modification

I

on the static longitudinal aerodynamic and stability characteristics. The results do not reflect the changes in reference area which resulted from the modification. The flap alteration promoted a significant reduction in

I

maximum lift, increased drag-due-to-lift; a._mall _, stable shift in longitu- dinal stability level; and reduced nose-up pitching moment due to sideslip

I

at moderate and high angles of attack. These results can be attributed to earlier breakdown of the wing leading-edge vortex system.

I

Figure 168 shows the effects of removing the inboard flap area on the lateral-directional stability characteristics. The modified flap promotes

I

lower levels of lateral stability at low and moderate angles of attack which is due to the weakened primary vortex system. At high angles of attack,

I

lateral stability is improved, although the variation of C_# with a is

I

oscillatory. Consistent with the part-span flap flow-field discussed in a previous section, a two-vortex system is developed. Since wing flow separa- tion is more pronounced at high angles of attack on the wing with the

I

part-span flap, there is less potential for significant vortex breakdown asymmetries in sideslip. Consequently, the lateral stability characterist-

I

ics are improved at high attitudes in comparison to results obtained with the unmodified, full-span flap. Directional stability is also improved above

I

25 ° angle of attack with the part-span flap.

I

The second flap modification was tested in an attempt to minimize the lift loss at high a's while retaining the increased roll stability associa- ted with the part-span flap. Relative to the wing with unmodified flap,

I

this configuration exhibits slightly reduced maximum lift, greater induced- drag, increased nose-up pitching moments with little change in longitudinal

i

stability, and more nose-up pitching moment due to sideslip, as depicted in Figures 169 and 170. Figure 171 shows the lateral-directional stability

I

characteristics. Evidently, this flap modification does not significantly alter the vortex flow field since the lateral stability characteristics are

I

similar for both flap configurations. At high angles of attack, the vortices that are shed from the two flap segments merge into a single,

I

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I

I

concentrated vortex system. Inherent in the latter is pronounced vortex

I

breakdown asymmetry in sideslip. The modified flap does, however, result in less directional stability at low-to-moderate angles of attack and reduced instability at high angles.

I

A systematic study of the effects of vortex flap apex geometry on the

I

longitudinal and lateral-directional aerodynamic and stability characteris- tics of fighter wings was performed in a parallel investigation under Air

l

Force sponsorship. Vortex flaps featuring curved leading-edges with zero- and finite-chord near the apex were tested along with vortex-flapped con- l, figurations with straight leading-edges. In all cases, the experimental data indicate that a curved apex with finite chord alleviates flow separ- ation from the deflected flap hingeline. However, as shown in Figure 172, a

I

straight leading-edge yields higher maximum lift, less drag at high lift, and reduced pitch instability. Figure 173 shows that both the straight and

I

curved leading-edge geometries provide lateral stability throughout the angle of attack range.

I

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Flap Deflection Effects A detailed study was conducted to assess the effects of leading- and

I

trailing-edge flap deflection angles on the 50-, 60-, and 65-degree cropped delta wings and of trailing-edge flap deflection angles on the 70/50-degree

I

cranked wing. The wing-flap geometries are shown in Figure 174. Full-span leading-edge vortex flap deflection angles of-30, O, 30, 45, and 60 degrees

I

were tested along with part-span trailing-edge flap deflection angles of O, 15, and 30 degrees. Differential deflections of trailing-edge flaps and

i

inverted leading-edge flaps for roll control were examined on the 60-degree cropped delta wing. It should be kept in mind that the vortex flaps were designed for a deflection angle of 30 ° .

i

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II

Leading Edge Flap Deflection Effects

il

The effect of leading-edge vortex flap deflection angles on the static longitudinal aerodynamic and stability characteristics are shown in I.

Figures 175-180. Increased flap deflection generally results in increased lift loss at low-to-moderate angles of attack, higher maximum lift, improved

I

drag polar, and reduced pitch instability. Nose-down pitching moment increments increase from _n = 30° to _n = 45°' but decrease from _n = 45° to _n = 60°"

I

As flap deflection angle is increased from 0 °, reductions in drag-due-to- lift are delayed to higher lift coefficients. In addition, nose-up pitching

i

moment due to sideslip is considerably reduced as flap deflection angle is increased. The data show that the effects are greatest on the 50-degree

I

cropped delta wing.

Upper surface flow patterns show that the leading-edge flap deflection li

delays the migration of the vortex off the flap to higher angles of attack.

However, when the vortex flap deflection is greater than 45 degrees, flow

II

separation and vortex formation from the hingeline occur. This flow situa- tion accounts for the small changes in the pitching moment characteristics

I

between flap deflection angles of 45 and 60 degrees.

I

The effects of leading-edge vortex flap deflection angle on the lateral-directional stability characteristics are shown in Figures 181 -

183. The 50-degree cropped delta wing exhibits significant reduction in I

lateral-directional stability at high angles of attack due to increased leading-edge flap deflection angle. The lateral-directional stability of

I

the 60- and 65-degree cropped delta wing were not significantly affected by increased leading-edge flap deflection angle. Increased flap deflection angle results in larger unstable increments in lateral stability at low-to- moderate angles of attack and delays the unstable C_. break to higher angles

i

of attack. However, leading-edge flap deflections of 45 and 60 degrees cause large fluctuations in lateral stability at moderate-to-high angles of attack due to the unsteady flow separation from the flap hingeline.

I

!

II

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I

The flap deflection angles does not significantly affect the direction-

I al stability derivative at low angles of attack. However, increased flap

deflection generally reduces the onset angle of attack for yaw instability with subsequent greater instability at high angles of attack.

I

Inverted Leading-Edge Vortex Flap Effects

I

The effects of a 30-degree upward deflection on the vortex flaps on the i longitudinal aerodynamic and stability characteristics of the 60- and 65-degree cropped delta wing are shown in Figures 184-187. The inverted flaps promote large lift increments at typical approach angles of attack,

I

primarily due to increased vortex strength. Inverted deflection decreases lift at higher angles of attack due to earlier vortex bursting, increases

I

drag as a result of an aft rotation of the vortex lift vector, and causes nose-up pitching moment increments as a result of the inverse camber. The

I

inverted flaps have a minimal effect on the longitudinal stability level.

However, the inverted flap configurations exhibit greater nose-up pitching

I

moment due to sideslip.

Figures 188-189 show the effects of inverted vortex flaps on the

I

lateral-directional stability characteristics. Except at very low angles of attack, the wings with upward deflected flaps develop much lower levels of

I

lateral stability. In addition, the inverted flaps promote an earlier unstable break in the lateral stability curve. These effects are due to the

I

more pronounced vortex breakdown asymmetries in sideslip. The inverted flap on the windward wing has an increased effective streamwise deflection which

i

results in increased vortex strength but, concurrently, earlier onset of vortex breakdown.

I

Trailing-Edge Flap Deflection Effects

l

Figures 190-221 illustrate the effects of trailing-edge flap deflec- tion on the static longitudinal aerodynamic and stability characteristics of

I

the 50-, 60-, and 65-degree cropped delta and 70/50-degree cranked wings.

I

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I

For a given leading-edge flap deflection angle, deflecting the trailing-edge

I

flap increases lift, reduces drag at moderate-to-high lift, and induces

large nose-downpitching momentincrements. In general, a pronounced

I

unstable break in the pitching momentcurve occurs at high lift with

trailing-edge flap deflected, due to earlier vortex breakdown. However, pitching momentdue to sideslip is greatly reduced.

I

Significant performance gains can be achieved by suitable combination

I

of vortex flap and trailing-edge flap deflections. Downward deflection

of the trailing-edge flap increases the upwashat the leading-edge at a

i

constant angle of attack, thereby increasing the suction pressures on the

deflected vortex flap. Thus, the deflected trailing-edge flap has a very

I

favorable influence on the vortex flap effectiveness.

Figures 222-225 present the effect of vortex flap deflection on

l

the longitudinal control derivative, C m . Vortex flap deflection

_f

I

generally reduces trailing-edge flap longitudinal control effectiveness

at low-to-moderate angles of attack. The inconsistent results associated

!

with a 60-degree vortex flap deflection are presumably due to the strong

vortex shed from the flap hingeline. At high angles of attack, pitch

I

control is increased on the cropped delta wings as a result of vortex flap

deflection. This effect decreases, however, with increasing wing sweep. In

all cases, as trailing-edge flap deflection is increased, flap longitudinal

i

control effectiveness diminishes. This is the result of trailing-edge flap

flow separation and earlier tip stall.

I

The effects of trailing-edge flap deflection angle on the lateral-

I

directional stability characteristics are shown in Figures 226-241. At

low angles of attack, trailing-edge flap deflection results in stable

I

increments in lateral stability which increase with increased deflection

angle. This effect is due to increased circulation about the windward wing

i

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I

t

and increased separation on the outward portion of the leeward wing. At

I

higher angles of attack, trailing-edge flap deflection generally results in large destabilizing effects on C_# and Cn# due to increased vortex breakdown

I

asymmetry in sideslip. The results obtained on the 50-degree cropped delta wing at angles of attack greater than 20 degrees should be interpreted with

I

the understanding that the unsteady flow separation from this wing at high angles of attack promotes large fluctuations in the lateral-directional

I

stability derivatives. This effect decreases rapidly with increased wing sweep, since a well-organized vortex flow can be maintained to higher angles of attack on the more slender planforms.

I

Vortex flap deflection enhances, and extends to higher angles of

i

attack, the dihedral effect due to trailing-edge flap deflection. In addition, the deflected vortex flap reduces the destabilizing effects

I

on C_# at high angles of attack arising from trailing-edge control surface deflection by limiting vortex breakdown asymmetry due to sideslip.

I

The lateral stability characteristics with vortex flaps deflected to

I

60 degrees warrant discussion. The highly-oscillatory variation of C_# with a is the result of the strong vortices that are shed from the leading-

I

edge flap hingeline region. Since there is no "aerodynamically-sharp" edge from which these vortices form, the vortex behavior is very sensitive to

I

model oscillation, tunnel turbulence, etc. Consequently, the hingeline vortex development and breakdown phenomena are unsteady.

I

A major shortcoming of inverted leading-edge flaps is illustrated in

I

Figure 234 corresponding to the 60-degree cropped delta wing. At typical approach angles of attack, the inverted leading-edge flap in combination with trailing-edge flap deflection promotes very large destabilizing effects

I

on lateral stability. This effect can be attributed to more severe vortex bur st asymmetry.

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Whenthe leading-edge flaps are undeflected, trailing-edge flap de-

I

flection typically has a favorable effect on the directional stability at

low-to-moderate angles of attack and delays the angle of attack at which Cn_

I

becomesnegative. Vortex flap deflection serves to diminish these effects.

I

Differential Flap Deflection Effects

I

The part-span trailing-edge flaps on the 60-degree cropped delta wing were deflected differentially such that the right flap was down 30 degrees

I

while the left flap remained undeflected. The leading-edge vortex flaps were symmetrically deflected to 30 degrees. Figures 242 and 243 show the

l

resulting rolling and yawing moment variations with angle of attack.

Results obtained at zero sideslip indicate that a differentially deflected

I

trailing-edge flap is a powerful roll control device at all angles of attack. Concurrent with the negative rolling moments are large proverse

I

yawing moment increments. Although one would expect positive yaw increments due to the increased induced drag on the right wing, this is not the case due to apparent wing leading-edge vortex-induced sidewash effect on the

I

centerline tail. Because of the proximity of the wing trailing-edge to the centerline vertical tail, the sidewash induces a positive sideforce, which

I

must be in the vicinity of the tail in order to promote a positive Cy, as shown in Figure 244. Vertical tail-off tests would be required to confirm

l

this hypothesis.

I

Figure 245 illustrates the roll control effectiveness of the differ- entially-deflected flaps. The configuration yields roll control effec- tivesess throughout the angle of attack range. The large yawing moments due

I

to trailing-edge control surface input are depicted in Figure 246.

I

Full-span leading-edge vortex flaps were deflected upward di ffer- entially on the 60-degree cropped delta wing. The right flap was deflected

I

up to 30 degrees and the left flap was undeflected. Trailing-edge flap deflection angle was 0 degrees. The resulting rolling and yawing moment

I

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I

variations with angle of attack are presented in Figures 247 and 248,

I

respectively. This configuration provides rolling moment increments at low angles of attack. At higher angles of attack, rolling moment increments in opposition to control surface input occur due to the earlier onset of vortex

l

burst asymmetries. This flow-field effect also results in destabilizing yawing moment increments.

I

The lateral-directional control derivatives are shown in Figures 249

I

and 250. The results show that differentially-inverted leading-edge flaps provide roll control at angles of attack less than 10 degrees. However, they exhibit undesireable reversal and roll control effectiveness.

I

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Vertical Tail Effects

i

The effects of outboard fins, fin deflection, and centerline tail deflection were examined in conjunction with the 70/50-degree cranked wing

I

with leading-edge vortex flaps deflected to 0 and 30 degrees. The outboard fins had the same tail volume as the centerline tail and were located at a

l non-dimensional span distance, y/s, of approximately 0.68. The principal

objective of this portion of the wind tunnel investigation was to obtain representative results showing the effects of outboard tail surfaces on the

I

leading-edge vortex behavior and longitudinal and lateral-directional stability and characteristics. No attempt was made to optimize the fin

I

location.

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Outboard Fin Effects

I

The effects of vortex flap deflection on the longitudinal aerodynamic and stability characteristics of the 70/50-degree cranked wing with outboard

l

fins are shown in Figures 251 and 252 for comparison to the vortex flap effects on the identical wing planform with centerline tail shown previously

I

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in Figures 22, 36, and 47. Consistent with previously discussed results,

I

vortex flap deflection reduces lift at a given angle of attack, reduces drag

at constant lift, has little effect on the pitching moment, and reduces the

It

pitching momentdue to sideslip. However, comparedto the flap deflection

effects on the configuration with a centerline tail, the configuration with

outboard fins yields reduced lift loss and better drag polar improvements.

II

Significantly, the outboard fin configuration continues to yield drag reduc-

tions throughout the lift coefficient range, while the centerline tail

I

configuration shows an increase in drag for lift coefficients above .9.

Pitching momentand longitudinal stability are similar for the two configu-

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rations although the outboard fin configuration does not show as much

improvement in the pitch-sideslip coupling as does the centerline tail

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configuration.

The effect of vortex flap deflection on the lateral-directional stabil-

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ity characteristics of the 70/50-degree cranked wing with outboard fins

are illustrated in Figure 9_53. The effects on the stability trends are

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similar to those on the 70/50-degree cranked wing with centerline tail

shown in Figure 58.

I1

A comparison of the outboard fin and centerline tail effects with

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vortex flaps deflected to 30 degrees on the static longitudinal aerodynamic

and stability characteristics are shown in Figures 254 and 255. Relative to

the centerline tail, the outboard fins result in significantly lower lift at

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angles of attack above 14 degrees, reduced maximum lift, higher drag-due-to-

lift at moderate and high lift, and more nose-up pitching moments. The

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reduced lift and increased drag are due to earlier vortex breakdownand a

lower effective aspect ratio associated with the addition of the outboard

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fins. In addition, the latter promote increased pitch instability and

nose-up pitching momentsdue to sideslip.

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These results can be better understood by examining the wing _pper

surface static pressure distributions. The location of the pressure ports

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on the 70/50-degree cranked wing are illustrated in Figure 256. Figure 257

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presents the effect of outboard fins on the wing upper surface static pressure coefficient. Results are shown at angles of attack of 16 and 24 degrees at two chordwise locations, x/c = .678 and x/c = .848. Sideslip

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angle is zero degrees. This discontinuity in the pressure distribution at the aft measurement station is due to the presence of the tail surface. At

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the forward measurement station, the pressure distributions with outboard fin off reveal a suction peak associated with the primary leading-edge

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vortex. The vortex has migrated off the inboard vortex flap segment and induces the inboard negative pressure peak at the aft measurement station.

The more pronounced suction levels near the leading-edge are due to a small, concentrated vortex on the deflected outboard flap surface. At the aft station, addition of the vertical fin promotes a large reduction in the suction pressures and an inboard displacement of the peak pressure associ- ated with the inboard vortex. These effects indicate vortex breakdown

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and a more inboard trajectory of the vortical flow. The concurrent reduc- tion in leading-edge upwash along the outboard wing panel results in lower

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suction pressures and an outboard shift in the peak pressure locations on the outer flap segment. An upstream effect due to the fins is evident

I in the pressure distributions at the forward station. Although the effects

are considerably less pronounced, the presence of the fins reduces the vortex-induced suction pressures. The surface pressure measurements are, therefore, consistent with the reduced lift, increased drag, and nose-up pitching moment increments associated with the fins.

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A comparison of the effects of outboard fins and centerline tail on the

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lateral-directional stability characteristics are shown in Figure 258. The fins reduce the dihedral effect at low angles of attack and significantly increase lateral stability at moderate and high angles. The reduced lateral

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stability at low angles of attack may be due simply to the lower position of the fin center pressure, and hence smaller moment arm, compared to that of

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the centerline tail. At higher attitudes where the leading-edge vortices are a dominant feature of the flow-field, the fins reduce and even eliminate

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vortex breakdownasymmetrydue to sideslip and promote corresponding

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large increases in lateral stability. In comparison to the configuration

with centerline tail the outboard fins reduce the directional stability and

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promote an earlier onset of yaw instability.

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The increased lateral stability due to addition of the outboard fins is

consistent with wing upper surface static pressure distributions. Figure

259 presents windward and leeward wing surface pressures at two measurement

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stations, x/c = .678 and .848, corresponding to angles of attack of 16 and

24 degrees and sideslip angles of 5 and 10 degrees. The results at the

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aft measurementstation reveal large reductions in the vortex-induced

suction pressures on the windward and leeward wings due to the presence of

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the fins. The effects, however, are generally more pronounced on the

leeward wing, particularly at large sideslip angle where inboard vortex

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impingement with the tail surface occurs. Under these conditions, .the

upstream influence of the fin on the wing surface pressures becomesmore

severe. In contrast, the windward leading-edge vortex is displaced away

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from the vertical fin as sideslip angle is increased, thereby mitigating the

tail influence on vortex stability. These factors result in enhanced roll

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stability at high angles of attack.

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Figure 260 compares the variations of rolling momentand lift co-

efficient with sideslip on the 70/50-degree cranked wing with outboard

fins and centerline tail. Consistent with the variations with sideslip of

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the peak suction pressure magnitude and location, the configuration with

outboard fins exhibits large variations in rolling momentwith sideslip

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relative to the centerline tail at the higher angle of attack.

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Figure 261 compares the dynamic directional stability parameter of the

outboard fin and centerline tail configurations. Relative to the configura-

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tion with centerline tail, the outboard fins promote positive increments to

C n at moderate-to-high angles of attack. This can be attributed

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principally to the reduced vortex breakdown asymmetryand, hence, improved

C_ characteristics shown previously in Figure 258.

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An examination of the surface flow patterns established on the cranked wing and outboard fin combination yields further insight into the flow-

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field. Figure 262 illustrates the surface flow patterns on the 70/50-degree cranked wing with vortex flaps deflected to 30 degrees, and on the windward and leeward sides of both vertical fins at 16 degrees angle of attack and 5

I

degrees of sideslip. For reference, the flow patterns on the wing without fin are also shown. The fins provide a fence-like effect, limiting the

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vortex-induced spanwise flow. In addition, comparison of the secondary separation line locations on the wings with and without fins indicates that

I

the presence of the fins causes an inboard displacement of the leading-edge vortices, particularly on the leeward wing. On the downwind, or suction, side of the leeward fin the flow is separated from the wing-tail junction to

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the tail mid-span and attached near the tip. On the inboard side of the leeward and windward fins the surface flow patterns reveal the scrubbing

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action due to the wing vortex interacting with the lower portion of the fin in addition to a tail-generated vortex near the tip. On the outboard, or

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pressure, side of the windward tail the flow is generally streamwise, indicative of attached flow.

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Additional evidence that, relative to the centerline tail, the outboard fins are immersed in an unfavorable flow-field is provided by tail surface

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static pressure measurements. Figures 263 and 264 show the effect of sideslip angle on the pressure distributions on the outboard fin and center-

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line tail suction and pressure sides ata= 16 and 24 degrees. The net lifting pressures, CPL, on the outboard fins indicate the latter are

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relatively ineffective lift-generating surfaces. In contrast, the center- line tail is exposed to less adverse vortex-induced sidewash and maintains

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lift effectiveness to high angles of attack.

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Tail Deflection Effects The effects of a +10 degree centerline tail deflection on the longitu- dinal and lateral-directional stability characteristics of the 70/50-degree cranked wing with vortex flaps deflected to 30 degrees are illustrated in Figure 265 and 266. The tail deflection has no effect on the longitudinal stability but results in greater nose-up pitching moments due to sideslip at moderate-to-high angles of attack. The configuration with the deflected tail displays less lateral stability at low-to-moderate angles of attack although a pronounced unstable break in the C_# curve is delayed to higher angles of attack. There is no change in directional stability due to tail deflection.

The effects of symmetric deflection of the outboard fins to +10 and -10 degrees on the longitudinal and lateral-directional stability characteris- tics are depicted in Figures 267 and 268. The tail deflections have no significant effect on the longitudinal stability or on the longitudinal and lateral-directional coupling parameter. Both deflection angles have a destabilizing effect on C_ up to about 24 degrees angle of attack.

Directional stability is also reduced due to fin deflection at low-to- moderate angles of attack, although the configurations with deflected fins exhibit slightly less directional instability at high a's.

Figure 269 presents the control effectiveness of the outboard fins with symmetric deflections of-10 degrees at zero sideslip. The roll control effectiveness, C_ , is favorable throughout the angle of v attack range. The yaw control effectiveness, Cn , diminishes and becomes 5v unfavorable at high angles of attack.

A comparison of the control effectiveness of the centerline tail and outboard fins deflected 10 degrees is shown in Figure 270. Consistent with the higher levels of lateral stability associated with the outboard _ fins,

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this configuration exhibits greater roll control effectiveness, relative to

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the centerline tail configuration. And, despite the lower levels of direct- ional stability, the deflected outboard fins are also a more effective yaw

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control device than is the deflected centerline tail.

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CONCLUDING REMARKS

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A detailed low-speed wind tunnel investigation was conducted to deter-

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mine the effects of analytically- and empirically-designed leading-edge vortex flaps on the static longitudinal and lateral-directional aerodynamic, stability, and control characteristics of fighter wings having sweep angles

!

of 45 degrees to 76.5 degrees. A determination was made of the sensitivity of the configuration forces and moments, wing surface pressures, and flow-

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field to symmetric and differential vortex flap and trailing-edge flap deflection angles; vortex flap planform, apex geometry, and controller tabs;

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trailing-edge sweep angle; wing position on the fuselage; nose strakes; closely-coupled canards; and centerline and outboard tail deflection angles.

The wind tunnel tests were performed in the Northrop 21-by 30-inch facility a

using an existing generic fighter fuselage model at angles of attack up to 40 degrees, sideslip angles to + 10 degrees, and free-stream dynamic

!

pressure of approximately 20 psf (130/ft/sec).

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The Vortex Lattice Method-Suction Analogy design procedure yields full- span vortex flaps of approximately constant chord planform on the wings

a

of moderate sweep angle (45 degrees to 55 degrees) and of inverse taper planform on the more highly-swept wings. The curved leading edge in the

apex and tip regions is an outcome of the vortex flap design procedure and !

has little basis in the real flow. The vortex flap size increases with increased wing sweep, and varies from approximately 10-15 percent of the

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wing area on the moderately-swept planforms to nearly 30 percent on the most slender wings. The design procedure is a useful tool for the preliminary

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sizing of leading-edge vortex flaps. However, the design criteria of vortex flow along the entire spanwise length of the deflected flap with flow

B

reattachment at the hingeline can only be met for the given set of design conditions by the wings of lower leading-edge sweep. As sweep increases, the vortex flow cannot be maintained on the flap and migrates onto the upper

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surface beginning at the tip and moving inboard with increasing sweep.

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In general, the V.LM-designed full-span vortex flap results in a

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reduction in the lift coefficient at angles of attack below the planar wing stall, increased maximum lift, decreased drag-due-to-lift, nose-down

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pitching moment increments, reduced pitch instability, and less nose-up pitching moment due-to-sideslip. These effects are typically more pro- nounced as vortex flap deflection angle is increased. Performance improve-

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ments due to a 30-degree vortex flap deflection normal to the hingeline are greatest on the wings of moderate sweep angle. On the more highly-swept

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_ngs, the large vortex flap size promotes a greater reduction in potential flow lift which limits the maneuver performance enhancement. A vortex flap

D

on a wing of given sweep angle exhibits only limited ability to maintain a concentrated vortical flow along the entire length of the deflected surface.

At typical transonic sustained maneuver lift coefficients, the migration of the vortex off the flap is delayed by increased flap deflection angle.

However, the favorable suction effect on the forward-facing surface is

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offset to an extent by flow separation from the deflected flap hingeline.

Experimental observations indicate that the leading-edge flaps deflected to

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angles up to 60 degrees (measured normal to the hingeline) are dominated by vortex separation at moderate-to-high lift coefficients. The presence of a

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concentrated vortical flow on the deflected flap is easily discernible in upper surface flow patterns and static pressure distributions. However, the

!

vortex structure in the cross-plane is not well-defined by an existing laser lightsheet technique except at high angles of attack where the vortex is large and of increased strength.

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Vortex flap deflection typically reduces lateral stability at low

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angles of attack due to flow separation on the flap underside. At moderate and high angles of attack, the deflected flap reduces tip separation, delays

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the onset of vortex breakdown asymmetry, and reduces the burst asymmetry once it occurs. These effects serve to increase the dihedral effect at

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those angles of attack. Exceptions to this trend are the 70-degree cropped delta and 76.5/66.5-degree cranked wings which experience adverse effects on roll stability at all angles of attack. This anomaly in the lateral

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stability behavior is an apparent result of the suppression of the

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leading-edge vortices at low-to-moderate angles of attack and a significant

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decrease in vortex strength at high attitudes due to deflection of the large

control surfaces. The lateral stability characteristics of a wing of given

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sweepangle, within the range of wing planforms examined in the present

study, are generally unsteady when the vortex flaps are deflected to large

angles. This effect is due to the development of a strong, but unsteady, I

vortex from the flap hi ngeline.

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The effect of vortex flap deflection on the directional stability is

not large. The yaw instability exhibited by the planar wing at high angles

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of attack is adversely affected by vortex flap deflection. In general,

the deflected vortex flaps promote positive increments to the dynamic

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directional stability parameter, C n , due to the increased dihedral

#dyn

effect.

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The effects on the longitudinal characteristics of VLM-designed

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partial-span vortex flaps that extend from 25 percent of the exposed wing span to the wing tip are similar, but less pronounced. In comparison to

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results obtained with the full-span flaps, the partial-span geometries yield consistently lower maximum lift, higher induced drag, and more nose-up

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pitching moments in conjunction with the 65- and 70-degree cropped delta wings. The test data show that the part-span flap results in a two-vortex system of reduced stability at high angles of attack.

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Comparisons of the lateral-directional stability characteristics of the

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cropped delta and cranked wings with full- and part-span flaps reveal a configuration dependence. The 65-degree cropped delta wing with full-span

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flap exhibits significantly higher lateral stability. In contrast, the 70-degree cropped delta and 70/50-degree cranked wings develop higher roll

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stability in combination with the deflected partial-span flaps.

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The aerodynamic and stability characteristics are very sensitive to small changes in the flap apex geometr_. Streamwise cuts to the apex

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regions of the part-span flaps on the 65-degree cropped delta and 70/50- degree cranked wings promote earlier bursting of the vortex system with resultant lift loss, drag increase, and increased pitch instability. The

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effects on lateral stability are configuration-dependent. The cropped delta wing develops significantly higher levels of lateral stability while the

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cranked wing exhibits consistently lower but stable dihedral effect through- out the angle of attack range tested.

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Inverted leading-edge flap deflection results in large increases in

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lift, higher drag, nose-up pitching moment increments, and more nose-up pitching moment due-to-sideslip at typical approach angles of attack. The increased vortex breakdown asymmetry due to upward deflection of the vortex

II

flap results in large, unstable increments to roll stability.

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Roll control due to differential deflection of an inverted leading-edge flap is limited to low angles of attack due to vortex breakdown effects.

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Leading-edge controller tabs are an effective means of manipulating the

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vortex strength at a given angle of attack. Relative to the plain vortex flaps, the tabbed flaps result in higher lift and less drag at high lift.

The increased vortex size arising from the tab deflection, however, causes

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an earlier migration of the vortical flow off the deflected flap. A compar- ison of the lateral-directional stability characteristics of the wings with

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plain and tabbed vortex flaps indicate that the former exhibit more lateral stability at the low angle of attack range. At high attitudes, however, the

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configurations with tabbed flaps maintain lateral stability while those with plain flaps become unstable.

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Empirically-designed vortex flaps of tapered planform yield increased

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lift, lower drag at moderate-to-high lift, and less pitch instability in comparison to the analytically-designed flaps of inverse taper planform.

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The inverse-tapered flap generally results in slightly higher levels of lateral stability due to its larger tip chord. The experimental results

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indicate that the flow situation corresponding to vortex separation every- where along the leading edge with primary reattachment at the hingeline is not necessarily "optimal ."

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Downward deflection of a part-span trailing-edge flap increases

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the vortex flap effectiveness due to increased upwash and the resulting higher suction pressures on the forward-facing surface. However, trailing-

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edge flap longitudinal control effectiveness decreases with increasing trailing-edge flap deflection angle due to flow separation on the trailing- edge flap. In general, trailing-edge flap deflection increases the lateral

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stability at low angles of attack, decreases the dihedral effect at higher angles, and promotes stable increments to the directional stability.

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In addition, vortex flap deflection enhances the trailing-edge flap pitch control effectiveness at high angles of attack.

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The aerodynamic and stability characteristics of a vortex-flapped wing

I

are sensitive to trailing-edge sweep angle. The results of the 65-degree cropped delta wing with vortex flap deflected to 30 degrees indicate that

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the cropped arrow wing (JLTE = +15 ° ) develops higher lift, lower drag, and reduced pitch instability compared to the cropped delta (A.TE = 0 °) and cropped diamond (ATE = -15 ° , -30 ° ) planforms. These results are due to

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the higher aspect ratio and larger ratio of vortex flap-to-wing area of the former wing. The cropped arrow planform exhibits better lateral stability

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characteristics at low and moderate angles of attack while the cropped diamond planforms maintain higher levels of lateral stability at higher

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attitudes.

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The results of testing the 65-degree cropped delta wing at three

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vertical positions show that the high- and low-wing positions result in increased lift, reduced drag at mid-to-high lift, and slight nose-up

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pitching moment increments with no change in longitudinal stability, relative to the baseline mid-wing position. Consistent with the classical fuselage crossflow effect on lateral stability, the low- and high-wing

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configurations exhibit lower and higher levels of lateral stability, respec- tively, than the mid-wing position. At moderate-to-high angles of attack,

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the high-wing configuration exhibits less directional stability while the low-wing configuration exhibits increased instability. Wing upper surface

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flow visualization revealed no discernible changes in the flow patterns due to wing height variations. In addition, upper surface static pressure

I distributions demonstrate that he magnitude and location fo the peak vortex-

induced suction pressures are relatively insensitive to wing vertical location on the fuselage.

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Strakes mounted at the nose maximum hal f-breadth have minimal effect on

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the total lift and drag characteristics of the 65-degree cropped delta and 70/50-degree cranked wings but induce nose-up pitching moment increments,

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increased pitch instability, and nose-up pitching moments due to sideslip.

The strakes result in increased lateral stability at moderate and high angles of attack, delay the onset of directional instability, and greatly

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reduce the directional instability at high angles of attack. The increased dihedral effect is due primarily to a favorable forebody-wing flow-field

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interaction. The improvements in directional stability are attributed to a direct strake vortex-induced suction effect on the forebody. The strakes

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also result in large, positive increments to Cn . Nose strake effects #dyn

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on vortex flap effectiveness are limited to pitching moment and directional stability characteristics: the strakes result in more nose-down pitching

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moment increments and reduced Cmi with flap deflection, and increase directional stability.

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