Table Of Contenthttps://ntrs.nasa.gov/search.jsp?R=19830010402 2019-04-05T20:40:24+00:00Z
NASA Technical Memorandum 84 5 8 2
Flap Effectiveness on Subsonic
Longitudinal Aerodynamic Characteristics
of a Modified Arrow Wing
P. Frank Quinto and John W. Paulson, Jr.
Lungley Research Center
Hampton, Virginia
Natimal Aeronautics
and Space Administration
S~krlUtfc-n b Technical
Informbclon Branch
SUMMARY
An investigation of the subsonic longitudinal aerodynamic characteristics of a
modified arrow-wing model was conducted in the Langley 4- by 7-Meter Tunnel. The
results of the investigation indicated that deflecting the leading edqe and trailing
edge in combination could promote an attached-flow condition at the wing leading
edge. Also, significant drag due to l i f t improvements resulted from combined leadinq-
and trailing-edqe flap deflections. The deflection of the trail ing-edge flaps
produced an increase in the upwash anqle, which caused the leading-edge vortex to
form a t a lower angle of attack. Leading-edge suction could be maximized over the
complete lift-coefficient range by scheduling a combination of leading- and trailing-
edge flap deflections.
INTRODUCTION
Highly maneuverable, supersonic-cruise fighter airplanes are under study by the
National Aeronautics and Space Administration (NASA), Department of Defense (DOD!,
and the aircraft industry. Tb aid in this study, the NASA Langley Research Center
has built a wind-tunnel model to investigate the subsonic aerodynamic characteristics
of a highly swept wing representative of such a fighter configuration. The planform
of this configuration is similar to that of the supersonic-cruise transport airplanes
that have been studied at Langley. Several investigations have been performed on the
subsonic aerodynamic characteristics of this transport configuration (see refs. 1
to 3i, and two of these investigations are on the effectiveness of leading- and
trailing-edge flaps. The results of these investigations helped in the design of
the present modified arrow wing, which w i l l be explained later in this report.
The arrow-wing model was tested in the Ianglev 4- by 7-Meter Tunnel to obtain
data on the effectiveness of the leading- and trailing-edge flap deflections in
achieving attached flow over the leading edge. This investigation addressed the
effectivensss of the flap deflections on the longitudinal aerody;~amic character-
istics, and the problem of trimming for this configuration was not considered. The
investigation included leading-edge deflections from O0 to 6Cb and trailing-edge
deflections from O0 to 30° a t angles of attack from -4O tc 24O at a Mach number
of 0.20.
SYMBOLS
All data have been rei11ced to coefficient form and are presented in the
s tabilit y-axis system. Computer symbols used are given in parentheses.
A aspect ratio
C (CD drag coefficient, Drag/qSref
D
induced drag coefficient
'D, i
=D, o drag coefficrent a t zero l i f t
ORIGINAL PAGE I3
OF POOR QUALITY
C~ (CL) l i f t coefficient, Lift./qsref
lift -curve slope, dCL/da
-
Cm (CM) pitching-moment coefficient, Pitching moment/qSrefc
-
c wing mean aerodynamic chord, 3.01 1 f t
L/D (L/D) lift -drag ratio, CL/CD
M ( MACH) Mach number
9 ( Q) f ree-s tream dynamic pressure, lbf/f t2
leading-edge suction parameter, percent (see eq. ( 3)
reference wing area, 10.422 ft2
u velocity in x-direction, ft/sec
w velocity in z-direction, f t/sec
coordinate pointing toward the nose and coordinate perpendicular
to the x coordinate, respectively
a (ALPHA) angle of attack, deg
flap deflection, normal to hinge line (positive, down;
negative, up), deg
Subscripts:
A attached flow
leading edge
separated flow
trailing edge
Abbreviations :
BL butt line
BS body station
PT ( PT) test-point number
MODEL DESCRIPTION AND TEST CONDITIONS
The leading edge of the modified arrow wing has an inboard sweep of 70° and an
outboard sweep of 48.8O. (See fig. 1.) lhe airfoil section is an NACA 0004, and the
wing has no twist. There are seven leading-edge flap segments per semispan, which
deflect normal to the leading edge. The maximum deflection of the inboard segment
is 20°. The maximum deflection of the next three segments increases a t lo0 intervals
(i.e ., .3 0°, 40°, and SO0 1; the last three outboard segments could be deflected up
to 60° An example of the leading-eP?e f lapdef lection terminology is 6LE = 40°,
whereby the actual leading-edge flap jegments are deflected from inboard to outboard
in the following manner: 20°, 30°, 40°, 40°, 40°, 40°, and 40°. We wing also has
three trailing-edge flap segments per semispan. The trailing-edge flap segments have
a wd.mum deflection of 30°. The leading- and trailing-edge flap segments can be
deflected independently of one another. (See table I for the wing geometry and both
fig. 2 and table I1 for the flap-deflection terminology used i n this report.)
A high-fineness-ratio fuselage was used to fair around the strain-gage balance
and other instrumentation. (See fig. 3.) The fuselage geometry is qiven in table I11
by the body station (BS) and its appropriate circular cross-sectional area. The
forces and moments were measured with a six-component strain-gage balance mounted
inside the model. The tests were conducted a: a dynamic pressure of 60 lbf/ft2,
which rt mlted in a Reynolds number of 1 .S x lo6 per foot at a corresponding Mach
number M of 0.20, and the angle of attack ranged from -4O to 24O.
DESIGN OF LEADING-EDGE FLAP DEFLECTION
In an effort to maintain attached flow on the leading edge of a highly swept
wing, a series of leading-edge flap& were installed which could be drooped a t various
angles as mentioned in the model-description section. lhe droop angles were chosen
so that the leading-edge flap would be approximately aligned with the flow upwash
angle (w/u + a) near the wing leading edge. By using the vortex-lattice method
NARUnE (North American Rockwell ' s Unified Vortex Lattice Extended Program described
in the theoretical-analysis section of ref. 4), an initial off-body velocity dis-
tribution was computed for the wing planform at a = lo0 and 6 = O0 and 30°
kE
with = OD. The leading-edge drmp was then set by the upwax angle indicated by
the initial off -body velocity distribution and used to predict the new off -body
velocity points with a = 1O 0 and = O0 and 30°. (The leading-edge deflection
was normal to the free stream during6 ~ is analysis, although the wind-tunnel model
had leading-edge deflections normal to the leading edge.) lhis procedure was
repeated until a converged solution (i.e., no change in velocity distribution or
leading-edge droop) was achieved.
The final upwash angles or leading-edge droop angles for both bTE = O0 and 30°
are shown in figure 4. These continuous droop distributions were then approximated
by the seven discrete flap elements on the wing leading edge. 'Ihe maximum deflection
of seven leading-edge flaps was chosen to approximate the maximum predicted upwash
angle a t the midpoint of the flap element.
PRESENTATION OF RESULTS
Table IV identifies the leading- and trailing-edge flap configurations of the
iging with the corresponding run number used in the wind-tunnel test8 the longitudinal
aerodynamic characteristics for the model are given in table V. Some longitudinal
aerodynamic results are presented as follows:
Figure
Effect of leading-ed ye deflections:
Figure
Effect of trailing-edge def lectionsr
DISCUSSION
Experimental Results
Effect of wing leading-edge flap deflections.- The longitudinal aerodynamic
characteristics at ,,6 = O0 with various leading-edge flap deflections are shown
i n figure 5. The lifimcoefficients at angles of attack below 6O are not appreciably
affected by the various leading-edge flap deflections. Also, at these small angles 1
of attack and low l i f t coefficients, the pitching moment is nearly zero, but it does I
have a slightly negative slope. In addition, the drag polar indicates a slight
increase in C with deflected leading edges. This is caused by the separation
D,o
of the lower surface flow at low angles of attack because of the leading-edge flap
deflection. At a moderate l i f t coefficient (0.3 < CL < O.6), the drag coefficient
decreases as the leading-edge deflections increase, which indicates ^that the
deflected leading edge is now carrying some leading-edge suction force.
A t higher angles of attack Ca > aO), the data indicate increases in l i f t coef-
ficient as the leading-edge vortex forms when the leading edge i-s undeflected. The
vortex-lift increase or increment diminishes as the leading-edge flap deflection
increases. A nose-up or unstable pitching moment begins a t a 16O. mis nose-up
moment is due to the onset of flow separation in the wing-tip region while the
inboard region is still maintaining attached flow. (See ref. 5.) Although the
pitching-moment coefficient is reduced as the leading-edge flap is deflected, the
pitching-moment slope still indicates an unstable pitching moment at a > 16O as
shown in figure 5.
The longitudinal aerodynamic characteristics a t ST= = 20° for various leading-
edge flap deflections are shown in figure 6. Comparison of the data of figure 5
( STE = 0°) with the data cf figure 6 (6 = 20°) shows that the 20° trailing-edge
TF
flap deflection increased the l i f t coefficient by about 0.3 and produced a nose-down
pitching moment of about 0.11 because of a trailing-edge deflection. Also, the
leading-edge vortex began a t a lower angle of attack for the undeflected leading edge
because of the increased upwash angle induced by t're trailing-edge deflection.
TO evaluate the effectiveness of the leading-edge deflections better, fiqures 5
and 6 also present two theoretical drag polars. These polars represent the planar-
wing minimum induced drag and the draq with full leading-edge separation with no
subsequent flow reattachment. This evaluation is similar to the ones performed in
references 1, 2, 3, and 6. The definitions of these drag polars are
for minimum induced drag or fully attached flow and
for full separated flow with no leading-edqe suction. Tbe value of drag coefficient
a t zero l i f t C for this configuration is 0.0021. The value of CD a is
D,o
obtained a t the CD intercept of a plot of 'P against C : tor an undeilected
wing. Elquations (1) and (2) are valid for a w ng with no twist and no camber and
are used as a quant~tativee valuator of the effectiveness of the leading-edge flap
deflections. 'Ihe leading-edge suction paraneter S is used as this evaluator and is
fined as
-
+ C tan a CD
C ~ , o L
- x 100
CL tan a
where CD and CL are measured l i f t and drag coefficients, respectively. For the
theoretical drag-polar calculations (CDIA and CD S), the value of a is replaced
is determined experimentaliy to be 0.037 (for the linear
where
region of CL against a for an undaflected wing), which agrees with theoretical
results from the NARWLE vortex-lattice program. 'Ihe leading-edqe suction parameter
is expressed in percent of leading-edge suction so that the minimum induced drag or
i full suction corresponds to S = 100 percent and full leading-edge separation with
no suction corresponds to S = 0 percent. (See ref. 6 for an explanation of the
leading-edge suction parameter.
Figure 10 presents the variation of leading-edqe suction parameter S with
CL for various leading-edge deflections wit!^ bTE = OO. At low lift coefficients
(0.2 < C C O.4), the suction parameter increases with increasing leading-edge
L
deflections until a deflection of bLE = 20° is reached. With higher deflections,
the suction parameter decreases because of overdeflection of the leading edge, which
causes f l w separation on the lower surface. nis increase in S is also true a t
-
higher lift coefficients (CL > 0.4). The highest suction parameter with an unc.
5
fleeted trailing edge is S = 66.6 percent a t bLE = 20° and = 0.31. This
value of lift coefficient may be small for approach or maneuver conditions; a more
appropriate value is about 0.6. At this l i f t coefficient, the maximum suction-
parameter value is S = 63 percent at bLE = 40°.
me
Effect of wing trailing-edge flap deflections.- longitudinal aerodynamic
characteristics with varying trailing-edge deflections are presented in figure 7 for
6LE
= 0°, in figure 8 for 6LE = 20°, and in figure 9 for = 60'. In these
bke
t ree configurations, the data indicate the expected increases in lift coefficient
and the pitching-moment coefficient becomes more negative with each trailing-edge
deflection. As the deflection of the trailing edge increases, a constant lift coef-
ficient can be achieved a t a lower angle of attack. For this situation, #here the
wing is operating near the zero suction curve = cL tan a), the angle of attack
(CD
is the dominant csrm and the induced drag coefficient is reduced. merefore, the
drag polar shows significant improvement a t higher l i f t coefficients as the flap
:lei lection is increased. Also, the lift coefficient increases proportionally to the
trailing-edge flap deflection for all three configurations until bTE = 30°. At this
flap deflection, the flow over the flap probably separates, thus causing a diminished
lift-coefficient increase. As mentioned in the previous section, the increase in
leading-edqe flap deflection causes the vortex l i f t to diminish for all trailing-edqe
flapdeflections (i.e., C is more linear).
La
The effect of the leading-edge flap deflection on the leading-edge suction
parameter for p. = lo0, 20°, and 30° is shown in figures 11, 12, and 13, respec-
tively. Ae in f gure 10 where 6Tf = 0°, the suction parameter increases as the
leading-edge flap deflections are ncrease-d until a maximum is reached; then, the
suction parameter decreases with further increases in leading-edge deflections. The
maximum suction-parameter value for bTE lo0 is S = 83.9 percent a t = 0.47
hE
with = 20°; for bTE = 20°, it is S = 86.6 percent at CL = 0.61 with
kE
= 30°; and for bTE = 30°, it is S = 83.6 percent a t
hE
= 30°. Also, note that as the trailing-edge flap
suction-parameter curves shift to higher l i f t coefficients and show a large leadinq-
edge suction value as shown in figures 10 and 11 to 13.
Effect of combined leading- and trailing-edge flap deflections.- 'he maximum
values of the suction parameter at various leading- and trailing-edge flap deflec-
tions are shown in figure 14. These values are from the peak numbers from figures 10
and 11 to 13. Fiqure 14 is, therefore, an envelope-type curve which presents the
maximum leading-edge suction for a combination of leading- and trailing-edge flap
deflections over a range of l i f t coefficients (i. e., higher leading-edge suction than
for fixed flap deflections). As l i f t coefficient increases, the maximum value of S
is about 86.6 percent a t E * 0.61 and decreases below 80 percent a t CL > 1.0,
which shows the difficulty n maintaining attached flow a t the high anqles of attack
required for high lift. For an approach or maneuvering configuration, C .= 0.6 and
k
the maximum value of S is about 86 percent a t bLE = 20° and bTE = 20 Fig-
ure 15 presents the effect of different leading- and trailing-edge flap combinations
on L/D at specific l i f t coefficients. The effect of the leading-edge flap deflec-
tion on L/D diminishes as l i f t coefficient increases. Also, as l i f t coefficient
increases, the peak L/D value shifts to a higher trailing-edge deflection. Compar-
ing the peak leading-edge suction values of figures 10 and 11 to 13 with the peak
values of L/D of figure 15 shows that the best leading-edge deflection is the same
in both peak S values and peak L/D values a t a specific l i f t coefficient. The
best trailing-edge flap deflection for the peak S values also occurs at the peak
L/D values.
A t the large leading-edge flap deflections ( bLE > 20° ), the leading edge becomes
stepped. The stepped secpei~tso f the leading edge cause some of the flow to sep-
arate. In order to determine the effect of the stepped leading edge on maintaining
attached flow, a comparison is presented in figure 16 of a faired and unfaired lead-
ing edge plotted against leading-edge suction parameter. The fairing causes the
leading edge to be a smooth flap rather than a stepped flap. mere is an increase
in the leading-edge suction parameter from the unfaired to the fafred ieadinq edge,
which indicates that the stepped leading edge is causing some flow separation; a t a
lift coefficient of 0.6, an improvement is obtained from S = 52.8 to 56.5 percent
by fairing the leading edge.
Theoretical Analysis
A preliminary theoretical analysis of the wing was conducted by using three
computer programs: (11 NARUVLE (see ref. 4) ; (2) a vortex-lattice program with
vortex-flaw computation using the leading-edge suction analogy (see .r efs. 7 and 8);
and (3 a surf ace-paneling program (PANAIR pilot code) (see ref. 9) Figure 17 com-
pares the experimental data with the results of the three theoretical programs for
the undeflected leadinq- and trailing-edqe case. The PANAIR pilot code and NARUVLE
program are attached inviscid-flow proqrams and do not account for the separated
vortex-flow condition of the wing. Only the lift-curve results from these two pro-
grams are presented. In the vortex-lattice programs of references 7 and 8, the
vortex flow is present to the tip of the wing. At the low angles of attack where
the flow is attached, these two programs agree with the experimental data.
The vortex-lattice program of references 7 and 8 accounts for the separated
leading-edge vortex flow of the entire wing and shows close agreement with experi-
mental lift data. Also, the potential-f low lift with high-angle-of -attack boundary
conditions is presented which causes the l i f t coefficient to be lower than NARWLE or
the PANAIR pilot code a t high angles of attack. Ihe theoretical drag polar from the
me
vortex-lattice program has the experimental C value added to it. slight
D, 0
difference in experimental and theoretical polars is due to a round leading edge
producing some leading-edge suction in the experimental data. A t a > 14O, the
theoretical results of this program do not correctly estimate the nose-up moment,
which can be due to the onset of flow separation in the tip area as discussed before.
Figure 18 shows the effect of trailing-edge flap deflections on l i f t as det~r-
mined by the NARUVLE program and by experiment. Ihe vortex-lattice program of refer-
ences 7 and 8 is not used in calculations of figure 18 because flap deflection cannot
be modeled easily. At low angles of attack (a < 4O), the theoretical results agree
with the experimental data up to 6TE = 15O. At higher trailing-edge deflections,
theory and experiment disagree because of the flow separation a t the trailing-edge
flap. At a = 8O and bTE = 0°, the two results, experimental and theoretical,
disagree because of the vortex flow caused by the undeflected leading edge which is
not modeled in the NARWLE program.
SUMMARY OF RESULTS
The results of an investigation of the leadinq- and trailing-edge flaps on an
arrow-wing model in the Langley 4- by 7-Meter mnnel are summarized as follows:
1. Def lectinq the leading-edge flap diminishes the vortex lift, which indicates
that the leading edge is approaching an attached-flow condition.
2. The drag due to lift is not significantly improved with the leading edge
deflected a t the desired l i f t coefficients (0.5 < cL < 0.9); however, the com-
CL
bined leading- and trailinq-edge flap deflections did result in significant improve-
ments in drag due to lift.
3. An increased leading-edge suction can be maintained throughout the lift-
coefficient range by scheduling both the lesding- and trailing-edge flap deflections
when compared with fixed flap deflections.
4. Close approximation of the longitudinal aerodynamic characteristics of
the undeflected arrow wing is obtained by using the vortex-lattice method of NASA
TN D-6142 and NASA TN D-7921 which accounts for vortex flow, except for the pitching
moment at high anqles of attack where the wing-tip region is most likely separated.
Langley Research Center
National Aeronautics and Space Administration
Hampton, VA 23665
January 7, 1983
REFERENCES
1. Coe, Paul L., Jr.; Thomas, James L.1 Huffman, Jarrett K.; Weston, Robert P.;
Schoonover, Ward E., . Jr.; and Gentry, Gar1 LO, Jr.: Overview of the LmWley
Subsonic Research Effort on SCR Confiqurations. Supersonic Cruise ~esearch
-
'79 Part 1, NASA CP-2108, 1980, pp. 73-33.
2. Coe, Paul L., Jr.; and Weston, Robert P.: Effects of Wing Leading-Edge Deflection
on Low-Speed Aerodynamic Characteristics of a Low-Aspect-Ratio Hiqhly Swept
Arrow-Wing Configuration. NASA TP-1434, 1979.
3. Coe, Paul L., Jr.; Huffman, Jarret-t K.1 and Fenbert, James W.: Leading-Edge
Deflection Optimization for a Highly Swept Arrow-Wing Configuration. NASA
TP-1777, 1980.
4. Tulinius, J.: Unified Subsonic, Transonic, and Supersonic NAR Vortex Lattice.
TFD-72-523, Los Angeles Div., North American Rockwell, Apr. 27, 1972.
5. Rao, Dhanvada M.; and Johnson, Thomas D., Jr.: Subsonic Pit ch-up Alleviation on a
74 Deg. Delta Wing. NASA CR-165749, 1981.
6. Henderson, William P.: Studies of Various Factors Affecting Drag Due to Lift at
Subsonic Speeds. NASA TN D-3584, 1966.
7. Margason, Richard J.; and Lamar, John E.: Vortex-Lattice FORTRAN Program for
Estimating Subsonic Aerodynamic Characteristics of Complex Planforms. NASA
TN D-6142, 1971.
13. Lamar, John E.; and Gloss, Blair B.: Subsonic Aerodynamic Characteristics of
Interactinq Lifting Surfaces With Separated Flow Around Sharp mges Predicted by
a Vortex-Lattice Method. NASA TN D-7921, 1975.
-
9. Moran, Jack; Tinoco, Edward N.; and Johnson, Forrester T.: User's Manual
Subsonic/Supersonic Advanced Panel Pilot Code. NASA CR-152047, 1978.
TABLE I.- GEOMETRY OF WING
...........................................................
Aspect ratio 1.639
........................................................
Reference area, ft2 10.422
...................................................................
Span, f t 4.133
.............................................................
Root chord* f t 4.7458
.........................................................
Tip chord, f t 0.9861
.......................................
Reference mean aerodynamic chord, f t 3.011
Leading-edge sweep, deg: ...............................................
At body station 29.672 in. .......................................... 70
At body station 86.620 in. 48.7858
.......................................... .........
Trailing-edge sweep, deg 24.5236
............................................................
Wing section NACA 0004
Description:modified arrow-wing model was conducted in the Langley 4- by 7-Meter has built a wind-tunnel model to investigate the subsonic aerodynamic
