Primary efforts directed toward drag reduction centered on the design of both supercritical and subcritical families of airfoils, the reduction of induced drag through the use of vortex diffusers, and the reduction of interference drag for executive-type aircraft

Kelly, T. C.

English

NASA Technical Reports Server

N76 11008
5.1 Some Methods for Reducing Wing Drag
and Wing-Nacelle Interference
Thomas C. Kelly
NASA Langley Research Center
]ntroduct ion
This paper summarizes the results of efforts directed toward drag reduction
which are underway within the Transonic Aerodynamics Branch at NASA-LRC. Three
areas of research which will be discussed are: (1) The development of both super-
critical and subcrltical families of airfoils (seer for examplet references 1 to 3);
(2) The development and application of vortex diffusers (ort more popularly_ winglets)
to reduce induced drag (reference 4); and (3) The application of supercrltical wings
to executive-type aircraft and the reduction of severe wing-pylon-nacelle interference
problems which were identified during these investigations (reference 5).
It should be noted that this is not a summary of the total NASA=Langley effort
devoted to ckag reduction, but rather a discussion of several selected areas which
would be of interest.
Work at Langley is continuing in several additional areas including the
reduction of turbulent skin friction through the use of compliant surfaces, reductions
in so-called "crud" drag through the application of various surface coverings, re-
ductions in induced drag using over the wing blowing (see reference 6), tip-mounted
engines (reference 7)_ favorable power interference effects, and the use of thick
supercritical wings to achieve higher aspect ratios. Finally, the possibility of obtain-
ing practical laminar flow control is also under consideration.
Airfoil Development
Supercritical airfoils - This new type transonic airfoil was developed by Dr.
R.T. Whitcomb at NASA-Langley about 10 years ago. At that time, theoretical
approaches for supercritical flows were non-exlstent and much of the early work in
developing the airfoils was intuitive in nature.
Figure 1 presents two-dlmensional wlnd tunnel results for several ten percent
thick airfoils at a section llft coefficient of 0.7. The conventional airfoil results
presented are for the NACA 64A-410. Comparison of the results for this airfoil with
those for the supercritical airfoils illustrates the significant gain in drag=rise Mach
number achieved by use of the supercritical section. The increase in drag-rlse Math
Preceding Page Blank
137
https://ntrs.nasa.gov/search.jsp?R=19760003920 2020-03-20T01:24:38+00:00Z
numberis approximately 0.1. These results also indicate the presence of &ag creep
which was characteristic of all of the early supercritlcal airfoils. The drag creep,
which results from increases in pressure drag associated with the onset of super-
critical flows on the airfoil upper surface, was noted in the early flight test results
for the F8 and T-2C research airplanes which employed supercritical wings having
these early-type airfoils.
Much of the recent work at Langley has been devoted to the el imlnatlon of
this undesirable drag creep, and the solid curve of Figure 1 shows the result of these
efforts. Refinements to the airfoil were involved primarily with changes which
resulted in a more favorable flow recompression over the forward upper surface and
the el imination of a region of flow overexpanslon near the three-quarter chord
location also on the upper surface. A slight loss in force-break or drag-divergence
Mach number is noted (about 0.01) as a result of slightly increased wave losses at the
higher Mach numbers, but this compromise is felt to be of little consequence relative
to the gains achieved in eliminating drag creep, it should be noted that, unlike the
early work, the shaping changes used in the design of the recent airfoil, were guided
by the use of the analytical program developed recently by Korn and others
(reference 8) in achieving desired pressure distributions for the various cases.
Figure 2 indicates the status of the current supercrltlcal airfoil development
program. As is evident, the main effort is ambitious, and covers a broad range of
design lift coefficients for airfoils ranging in thickness ratio from 6 percent to 22
percent. Solid symbols on the figure indicate airfoils which have been designed and
tested. Open symbols represent airfoils which are currently under design using the
analytical program noted earlier_ and the Crossed symbols indicate airfoils which
are considered to have important applications and which are planned for design in the
future.
Subcritlcal airfoils - The Langley work on supercrltical airfoils led to a renewed
interest on the port of NASA in developing airfoils designed for subcrltical speeds.
The first of these, designated the GA(W)-I, was a 17-percent airfoil designed
spec iflcally for the slngle-engine cl imb requirements for I ight tw in-eng ine a ircraft.
[nitial wlnd-tunnel results for this airfoil encouraged the design of several other
airfoils for use at subcrltical speeds.
Figure 3 presents a comparison of two-dimenslonal results obtained recently
for the 13-percent thick GA(W)-2 airfoil and the older NACA 651-213 airfoil.
Results for both airfoils were obtained in the Langley low turbulence pressure tunnel
using the narrow fixed transition strip technique. Assuming section lift coefficients
138
of 0.4 and 1.0 for the cruise and climb cases, respectively, reductions in section drag
coefficient on the order of 3 percent and 17 percent, respectively, are achieved for
the newer alrfoil. Significant gains at the higher lift coefficients are indicated.
Figure 4 corresponds to the earl ier supercritical family figure and indicates
the current status of the subcritical airfoil family development. All of the airfoils
developed thus far have been designed for a llft coefficient of 0.4, having thlckness
ratios of 13, 17, and 21 percent. An additional airfoil in this group is being designed
with a thickness ratio of 9 percent. Also in design are two 17-percent airfoils having
design lift coefficients of 0.2 and 0.7. Planned for future development are two
13-percent airfoils with design llft coefficients of 0 and 1.2.
Vortex Diffusers
A sizeable effort is currently underway at NASA-LRC directed toward the
reduction of induced drag. One phase of this effort, which has received wide notice,
involves the use of vortex diffusers, or wlnglets as they are more popularly called,
mounted at the wing tips. Unlike end plates, the vortex diffusers are designed with
the same careful attention to local flow conditions as would be utilized in the design
of the wing itself. The placement of the vortex diffuser within the rotational flow
field of the wing tip results in forward inclinations of the lift (or side force) Vectors
far the vortex diffusers, producing a thrust component which increases with increasing
lift. The action is analogous to the force which propels sailboats, of course.
Figure 5 shows the geometric characteristics of the semispan model used for the
vortex diffuser development. The wing planform represents an early version of a
wide-body transport configuration. The two tip configurations tested, shown in the
right side of the figure, represent the basic tip configuration and the vortex diffuser
configuration which was tested assuming the "soft tip" portion of the basic wing panel
could be removed. This resulted in a reduction in the basic panel span of about
2.5 percent. The vortex diffuser geometry is shown on the left side of the figure.
The upper vortex diffuser span is about equal to the basic wing tip chord; the leading
edge sweep is equal to the wing leading edge sweep and the vortex diffuser root chord
extends over the aft 60 percent of the wing tip chord. This particular arrangement
was selected in order to avoid superimposing the high local velocities occurring on
the wing upper surface and on the vortex diffuser upper surface, which faces inboard.
This upper diffuser is canted outward about 18° from the vertical. The lower diffuser
is reduced considerably in span to provide ground clearanoe, extends over the forward
40 percent of the wing tip chord, and is canted outboard about 36 ° . The upper and
lower vortex diffusers were separated to avoid mutual interference effects.
139
Figure 6 presents incremental drag results associated with the vortex
diffusers as a function of llft coefficient for three Math numbers. Incremental drag
is defined as the drag coefficient for diffusers on minus the drag coefficient for
diffusers off so that negative values of this parameter represent gains or, thrust.
Near zero lift, a net penalty results as would be expected. At a lift
coefficient of about 0.26 the diffusers are carried with no penalty, and above this
lift coefficient, favorable effects are obtained which increase with increasing lift.
For the case presented, the cruise lift coefficient is about 0.53 at a Mach number
of 0.80, resulting in a gain of about 15 drag counts ( C D = 0.0015). In full-scale
terms this would represent an increase in L/D of about 5 percent, which, of course, is
significant.
Figure 7 indicates the effect of the vortex diffusers on the wing pitching-
moment and root-bending-moment coefficients. These results are for the cruise
Math number of 0.80, and show relatively small effects of the diffusers, for example,
a two percent increase in root bending moment at the highest lift coefficient tested.
Also, early wind tunnel flutter tests have indicated relatively small reductions in
flutter dynamic pressure resulting from addition of the diffusers. It was also concluded
that these effects were associated with structural characteristics rather than any
unsteady aerodynamlc interaction.
Figure 8 illustrates one proposed application of vortex diffusers. NASA is
currently involved in a joint program with the USAF to determine the possibility of
adding vortex diffusers to both the C-141 and the KC-135. Tests of both configurations
are scheduled for this fall in the 8-foot Transonic Pressure Tunnel. This model photo-
graph depicts vortex diffusers installed on the C-141.
Wing-Nacelle Interference
In an effort to provide access to supercritical wing technology for the
general aviation manufacturers, NASA has entered into several cooperative endeavor
agreements with members of the industry whereby NASA provides expertise in the
areas of aerodynamic design and application of supercrltical wings to this class of
aircraft and also provides limited wind tunnel testing for configurations which are under
design. The remainder of this discussion will present some results obtained recently
which relate to the problem of wlng-nacelle interference wh ich occurs at high
subsonic speeds and which is characteristic of configurations where fuselage-mounted
engines overhang the wing rearward upper surface. A representative configuration is
shown in Figure 9. This is a photograph of a one-nlnth scale model of an
140
executive-type aircraft under test in the Langley eight-foot transonic pressure
tunnel. This part icular investigation was conducted to determine the aerodynamic
feasib;I ;ty of replacing the original wing with one having supercritical sections of
increased thickness, the thickness ratio being increased on the average from about
.09 to .12. Because the wing was intended to be retrofitted, other configuration
changes were to be kept to a minimum, therefore most of the tailoring or "tuning"
changes were associated with the wing itself although some changes to the pylons
were also made.
Figure 10 illustrates the severity of the interference problem which was found
to exist between the wing and the engine-pylon arrangement for the modified con-
figuration. These results are for a lift coefficient of 0.25 and are presented as an
incremental drag coefficient versus Mach number, where M -- 0.60 is used as a
reference drag level for each configuration. It should be emphasized that all of the
results presented are for the configuration with a supercrltical wing. No comparisons
are presented with the basic wing.
The data for the initial supercritical wing configuration, shown by the solid
curve, indicate noticeable drag creep and an early drag rise. Reasons for these
effects will be discussed later. Removal of the engines and pylons resulted in sig-
nificant reductions in drag-creep, and the force-break or drag-rise Mach number
was increased by about 0.02 to 0.03 as indicated by the dashed curve. Modifications
to the wing root section, the addition of a wing glove and some reshaping of the
pylon provided the results indicated by the long-short dashed curve. As can be
seen, the drag creep was reduced considerably (about 75 percent at a Mach number
of 0.80).
Figure 11 presents wing upper-surface oil flow photographs for a Mach number
of 0.825 and a lift coefficient of about 0.35. The photograph on the left of the
figure is for the initial configuration with nacelles and pylons and shows dramatically,
the effect of the nacelle and pylon presence in forcing the uppers urface shock wave
forward on the inboard wing region. The presence of a second wave which originates
in the channel formed by the upper surface of the wing and the nacelle-pylon combina-
tion is also apparent. Examination of local pressure distributions for this case indicates
the second wave and the adverse pressure gradients associated with it caused extensive
separation in the "channel" region. The center photograph, for the case of the
nacelles and pylons removed, shows what could be termed an expected supercritical
wave location for this thickness ratio on the upper surface with little or no separated
flow in evidence.
141
Thefinal photographof Figure 11 illustratesthe uppersurfacecondition for
the tunedconflguratlon. Themainwave appearssomewhatweaker compared to that
for the initial configuration, and no evidence of the second wave is seen. Modifica-
tions to the configuration were accomplished through a number of steps. Figure 12
shows a comparison of the initial wing root and pylon lines with those for the final
configuration tested. Most of the changes noted were made to eliminate what was
essentially a converging-diverging channel formed by the wing-pylon-nacelle
combination. Initially, material was removed from the aft region of the wing upper
surface. Then, in order to reduce the upper surface velocities entering the channel,
a glove was added and the forward portion of the resulting airfoil was increased in
thickness. Finally, the lower surface of the pylon was thickened in order to provide
for a relatively constant area in the channel. All of the noted changes represented
an application of the area rule in a local sense, since the induced velocities
approached and exceeded sonic values in the region.
The results just disoussed were obtained using an executlve-type aircraft
model which employed turbojet engines. It might be expected, therefore, that
replacement of the turbojet engines with the newer and larger turbofan engines would
serve to aggravate the interference problem which was noted and some recent wind
tunnel results indicate this to be the case.
Figure 13 presents results for an executive-type aircraft very similar to the
one previously discussed. Again, results are presented in the form of an incremental
drag coefficient versus Mach number; however, the reference Mach number in this
figure is 0.50. Because an interference problem was anticipated, provision was made
on the model to allow translation of the engine-pylon combination to several longi-
tudinal locations. The results given in Figure 13 are for the proposed initial location
(identified in Figure 13 as the product|on location), the most rearward location
possible, which corresponded to a full-scale rearward shift of 18 inches and for a
final "tuned" configuration with the engines aft. For this span location, and neglecting
the glove, the leading edge of the nacelle moved from about the 55-percent chord
location to the 70-percent location. This relocation was made in three steps of 6
inches each, full scale, and the results indicate that further gains could be achieved
by additional rearward movement of the nacelle. Obviously, however, airplane
balance problems would impose some practical rearward limit to the nacelle reloca-
tion. As with the earlier results, noticeable additional gains were made through
so-called "tuning" which, in this instance, involved primarily changes in the
inboard airfoil shape.
142
Concluding Remarks
Efforts underway within the Transonic Aerodynamics Branch at NASA-LRC
have been directed toward drag reduction in several areas. Primary efforts have been
involved with the design of both supercritical and subcritlcal families of airfoils, the
reduction of induced drag through the use of vortex diffusers, and the reduction of
interference drag for executive-type aircraft•
The results of many of these efforts are felt to be applicable to the design of
general aviation aircraft.
References
•
o
o
•
o
a
7.
o
Whltcomb, R.T., "Review of NASA Supercritical Airfoils," ICAS Paper No.
74-10, The Ninth Congress of the International Council of the Aeronautical
Sciences, August 1974.
McGhee, R.J., and Beasley, W.D., "Low-Speed Aerodynamlc Characteristics
of a 17-Percent-Thlck Airfoil Section Designed for General Aviation Appl!ca-
tionst NASA TN D-7428, 1973.
McGhee, R.J., Beasley, W.D., and Somers, D.M., "Low-Speed Aero-
dynamic Characteristics of a 13-Percent-Thlck Airfoil Section Designed
for General Aviation Applications, NASA TM X-72697.
Flechner, S.G., and Jacobs, P.F., "The Effects of Vortex Diffusers (Wing-
lets) on a Semispan Model of a Representative Wide-Body Transport," Proposed
NASA TN D.
Bartlett, D.W., "Application of a Supercritical Wing to an Executive-Type
Jet Transport Configuration," Proposed NASA TM X. (L-9939).
Putnam, L.W., "Exploratory Investigation at Mach Numbers from 0.40 to
0•95 of the Effects of Jets Blown Over a Wing," NASA TN D07367, 1973.
Patterson, J.D., and Flechner, S.G., "An Exploratory Wind-Tunnel Investiga-
tion of the Wake Effect of a Panel Tip-Mounted Fan-Jet Engine on the Lift-
Induced Vortex, i. NASA TN D-5729, 1970.
Bauer, F., Garabedlan, P., Korn, D., and Jameson, A., "Supercrltical
Wing Sections II," Lecture Notes in Economics and Mathematical Systems,
Springer-Verlag, c. 1975.
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Some methods for reducing wing drag and wing-Nacelle interference

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