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NASA
Technical
Paper
2273

AVSCOM
Technical
Report
83-A-17

February 1984

NASA

ENGINEERING

Airfoil Interaction With
an Impinging Vortex

K. W. McAlister
and C. Tung

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UNIVERSITY OF ILLINOIS-URBANA

3 0112 106710061

NASA
Technical
Paper
2273

AVSCOM
Technical
Report
83-A-17

1984

WASA

National Aeronautics
and Space Administration

Scientific and Technical
Information Branch

Airfoil Interaction With
an Impinging Vortex

K. W. McAlister
and C. Tung

Aeromechanics Laboratory

USAAVSCOM Research and Technology Laboratories

Ames Research Center

Moffett Field, California

SYMBOLS

C chord of downstream airfoil, m
Cj drag coefficient

Cn lift coefficient

C quarter-chord pitching-moment coefficient

c chord of generator airfoil, m

Re Reynolds number, U^C/v

r radial distance from the vortex center, m

[/qq free-stream velocity, m/sec

w circumferential -velocity component, m/sec

a airfoil incidence, deg

a generator incidence, deg

T circulation

v kinematic viscosity, m2 /sec

Digitized by the Internet Archive

in 2013

http://archive.org/details/airfoilinteractiOOmcal

SUMMARY

The tip of a finite-span airfoil was used to generate a streamwise vortical flow, the
strength of which could be varied by changing the incidence of the airfoil. The vortex that
was generated traveled downstream and interacted with a second airfoil on which measure-
ments of lift, drag, and pitching moment were made. The flow field, including the vortex
core, was visualized in order to study the structural alterations to the vortex resulting from
various levels of encounter with the downstream airfoil. These observations were also used
to evaluate the accuracy of a theoretical model.

1. INTRODUCTION

The vortices that are generated by missiles, canards,
wings, and rotor-blade tips often have a detrimental effect
on the flow fields of other control or lifting surfaces. One of
the most elementary models of this type of flow interaction
is provided by the passage of a streamwise vortex near a
downstream lifting airfoil. For an accurate calculation of this
flow field, it is necessary to correctly account for (1) the
time-varying viscous structure of the vortex; (2) the three-
dimensional viscous flow over the airfoil, including the shed-
ding of its own wake; and (3) the nonlinear path of the
vortex resulting from its interaction with the airfoil. From
the experimenters' point of view, the challenge is (1) to pro-
duce a fully developed, steady, and well-defined vortex in
the flow, without the attendant wake of the generator,

(2) to correctly scale the vortex-airfoil interaction, and

(3) to provide suitable measurements in sufficient detail to
meet the level of evaluation required.

The mathematical model for the impinging vortex has
ranged in complexity from that of an in viscid -line vortex
fixed along a rectilinear path, to a viscous-core vortex devel-
oping along an unprescribed path. Similarly, the mathemati-
cal model for the interacting airfoil has evolved from a simple
lifting-line theory to a dense vortex-lattice representation
(refs. 1-3). Numerous experiments have been performed to
assess the value of various combinations of these computa-
tional models, as well as to define the flow field and resultant
loads on the airfoil during the interaction. These studies have
shown that when details of the flow are required (such as air-
foil pressure distribution) during a close vortex encounter
(roughly within one core diameter), only the most compre-
hensive models are capable of providing calculations with
acceptable accuracy. In those cases in which the vortex inter-
action is severe enough to cause separation on the airfoil, the
choice of models must be narrowed to the few that include
the boundary layer. Furthermore, the boundary -layer model

must be three dimensional to account for the strong spanwise
flow component caused by the interaction (ref. 4). Recogni-
tion of the boundary layer is an important factor in deter-
mining the full effect of the vortex-airfoil interaction since
vortex-induced separation on the airfoil has been found to
substantially limit the extent of the induced loads (ref. 5).
Only recently have codes become available that are capable
of treating the vortex interaction problem where flow separa-
tion is present (ref. 6), and the results from one of these will
be examined in light of the present experiment.

Although many noteworthy vortex interaction studies
have preceded this investigation, some aspects of the problem
have not been sufficiently addressed and therefore remain in
question. Specifically, these questions concern the alterations
to both the trajectory and stability of the vortex, as well as
the overall performance of the airfoil resulting from the
interaction. This subject can be most simply addressed by
considering the case for a streamwise-oriented vortex encoun-
tering a two-dimensional lifting airfoil. Those questions per-
taining to the vortex are (1) Does the path of the vortex
essentially conform to the streamline pattern existing for the
airfoil alone? (2) To what extent does the strength of the
vortex influence its trajectory? and (3) Is proximity to the
airfoil sufficient to cause an appreciable diffusion or break-
down of the vortex? Those questions regarding airfoil per-
formance are (1) How does the presence of a nearby vortex
(either passing above or below the airfoil) affect the airfoil
stall? and (2) To what extent are the total pre-stall loads on
the airfoil affected by a direct vortex impingement? These
questions were to be addressed in the present experiment by
visualizing the vortex and the airfoil boundary layer, along
with direct measurements of airfoil lift, drag, and pitching
moment.

In addition to obtaining certain physical insights into the
subject of vortex-airfoil interactions, there was an interest in
comparing the results of the experiment with the calculations
of a promising mathematical model. This comparison would
not only provide an opportunity to evaluate the accuracy of

the model, but would also form the basis on which any
refinements to the model are made.

The authors would like to acknowledge and express their
appreciation to Rabindra Mehta, T. T. Lim, and Raymond
Pi/.iali, who reviewed the original manuscript. They provided
valuable challenges to various technical issues raised by the
authors, and in so doing, contributed greatly to the readabil-
ity and accuracy of the final report. The authors would also
like to thank Brian Maskew (Analytical Methods, Inc.) for
contributing the theoretical model, for supporting the com-
parison with the experimental results in an unbiased manner,
and for so kindly providing counsel whenever it was required.

2. DESCRIPTION OF THE EXPERIMENT

This study was conducted in the 4000-liter, closed-
circuit water tunnel facility at the Aeromechanics Labora-
tory, Ames Research Center (fig. 1). This was a particularly
suitable facility for this investigation because of the ease of
obtaining definitive visualizations of the vortex and the
advantage of examining on-line the resultant loads on the
airfoil during the interactions. The technique for visualizing
the flow was based on the generation of minute hydrogen
bubbles through electrolysis of a weak solution of sodium
sulfate and water. Loads were measured directly by an exter-
nal apparatus that served as both support and balance for the
airfoil.

The airfoil selected for this study was a NACA 0012
having a two-dimensional planform of 10 cm (chord) by
21 cm (span). The test section measures 31 cm (height) by
21 cm (width), and the airfoil was positioned so that it
spanned the width of the section to within 0.015 cm on
either side. The airfoil was cast of an electrically nonconduct-
ing fiber resin, with platinum electrodes placed at nine chord-
wise locations along the upper surface. The bubbles that were
generated at these electrodes were transported downstream
by the fluid in the boundary layer and wake, thus enabling
the thickness and eventual separation of the boundary layer
to be observed.

The vortex was generated by placing a semispan airfoil
at incidence in the free stream ahead of the NACA 0012 air-
foil. The vortex generator was a NACA 0015 airfoil with a
rectangular planform and a 5-cm chord (fig. 2). Two vortex
generators were constructed from an electrically nonconduct-
ing fiber resin. When installed, in turn, on the upper test sec-
tion wall (fig. 3), the tip of one generator would extend to
the centerline of the tunnel and therefore be on line with the
pitch axis of the downstream airfoil (generator aspect ratio
of 3); the tip of the other generator would be 0.5 c above
the downstream airfoil (generator aspect ratio of 2). Two
electrodes were placed on each vortex generator. One of the
electrodes was located on the pressure side of the generator;
it extended over 80% of the chord in a streamwise direction

and was inboard from the tip a distance of 0.1 c. This elec-
trode was used to visualize the tip vortex. By generating
bubbles on the pressure side and allowing them to be
advected around the tip to the suction side, the authors
believe that a more accurate picture of the coalescing and
shedding behavior of the tip-vortex core is obtained. The
second electrode was located on the suction side of the gen-
erator, extended over 1 .3 cm in a spanwise direction, and was
upstream from the trailing edge a distance of 0.2 c. This elec-
trode was used to monitor flow separation on the generator.
A third electrode was attached to the tip of the generator at
the quarter-chord location, and was stretched across the flow
to a connection point on the lower test-section window. The
purpose of this electrode was to visualize the helical structure
of the vortex outside of the core region. The pitch axes of
both the generator and the airfoil were located at their
respective quarter-chords, and a distance of four generator-
chord lengths separated the two axes (fig. 4). This arrange-
ment provided a vortex maturation distance of 2.75 c from
the trailing edge of the generator to the leading edge of the
airfoil.

The spar of the airfoil extended through the test-section
windows and was supported by lift and drag transducers on
both sides (fig. 5). One end of the spar was adjoined to an
instrumented drive shaft through a torsionally stiff coupling
so that airfoil incidence could be set and the pitching
moment measured. Static frictional moments imparted by
the support bearings and seals were also measured and later
treated as load tares. Only quantities relating to the airfoil
were electrically instrumented: incidence, lift (both sides),
drag (both sides), total pitching moment, and the bearing and
seal moments (both sides). After amplification, the signals
were either appropriately summed (i.e., total pitching
moment minus both frictional moments) and displayed on
local monitors or they were transmitted to a remote data
acquisition system where they were digitized, averaged, and
stored for later processing. It is estimated that both airfoil
and generator incidence were set to an accuracy of 0.2
during the test. Lift and drag measurements are considered to
be accurate to 0.01 N and the pitching moments to
0.002 N-m.

The bubbles were illuminated by a sheet of light (about
5 cm wide) directed through the upper test-section window
and covering a length of 30 cm in the free-stream direction
(fig. 6). Both continuous and flash sources of light were pro-
duced over this length. The continuous source of light was
provided by a single 1000-W quartz-halogen lamp; the lamp
was used for general viewing, as well as for long-duration
exposures (20 sec in this experiment). The flash source of
light was obtained from a 10,000-W xenon lamp that could
either be synchronized to the shutter of a high-speed camera
or operated in a single-flash mode with a view camera. A
second xenon lamp (not shown in fig. 6) was directed
upward through the lower test-section window to provide an
equal amount of illumination from below the airfoil.

The tunnel was operated at two fixed drive speeds dur-
ing this experiment. With the airfoil set at zero incidence, the
dynamic pressures for these two speeds were 0.10 lb/in.2 and
0.025 lb/in.2 ; they are equivalent to Reynolds numbers of
120,000 and 60,000, respectively, based on an airfoil chord
of 10 cm. Some reduction in tunnel speed is thought to have
occurred when the airfoil was stalled; however, no attempt
was made either to measure or account for this degradation.

The scope of the experiment was limited to discrete
values of incidence for the generator and airfoil. The airfoil
was placed at both positive and negative values of incidence,
and at angles ranging from 0° to beyond stall (in 1° incre-
ments). Three free-stream conditions ahead of the airfoil
were considered. First, a control case in which no vortex
generator was present. Second, a mild interaction case result-
ing from a short vortex generator (tip off centerline) being
placed in the stream at angles of 0°, 5°, and 10°. Third, a
severe interaction case resulting from a long vortex generator
(tip on centerline) being placed in the stream at angles of 0°,
5°, and 10°. Lift, drag, and pitching moment measurements
on the airfoil were made at Re = 120,000. Flow visualiza-
tions were made at both Re = 60,000 and Re = 120,000,
with corresponding velocities of 0.58 m/sec (1.9 ft/sec) and
1.16 m/sec (3.82 ft/sec).

3. DESCRIPTION OF THE THEORY

The particular theoretical model to be used for compari-
son with the experimental data is a panel method formula-
tion using Green's theorem. The code is capable of calculat-
ing the trajectory of the vortex, as well as the resulting loads
on the airfoil arising from the interaction. A detailed descrip-
tion of the method is given in reference 7; however, a brief
discussion will be presented here for convenience.

The surface of the wing is approximated by a set of flat
panels consisting of uniform sources and doublets. The
source strength of each panel is determined by the local
external Neumann boundary condition and the strength of
each doublet distribution is determined from a set of simul-
taneous linear equations explicitly specifying the internal
Dirichlet boundary condition of zero perturbation potential.
The wake generated by the flow over the airfoil is also repre-
sented by flat panels of uniform doublet singularities. All
wake panels along a streamwise column have the same doub-
let strength as determined by the zero-load condition at the
trailing edge heading that column. When the flow is separated
from the leading edge, the wake is enclosed by a pair of free-
shear surfaces, each having a doublet distribution of linear
strength in the streamwise direction and of constant strength
in the crossflow direction. The code also provides for a fully
coupled boundary-layer calculation in order to account for
the viscous-inviscid interaction.

4. DISCUSSION OF RESULTS

Flow Visualization at Re = 60,000

The tip of the vortex generator was located on the cen-
terline of the tunnel and was, therefore, geometrically on line
with the pitch axis of the downstream airfoil. The vortex
generator was set to three angles of incidence, a - 0°, 5°,
and 10°; and for each of these angles the downstream airfoil
was varied from -16° to +16° (figs. 7-9). By placing the
generator at 0° incidence, a control case (fig. 7) was estab-
lished against which the effects of the vortex on the stream-
lines around the airfoil could be evaluated. For brevity, the
upstream airfoil that was responsible for producing the tip
vortex will be referred to simply as the "generator" while the
downstream airfoil that interacted with the vortex will be
referred to as the "airfoil."

Rotating the generator to 5° incidence caused a weak
vortex to be produced (fig. 8). The hydrogen bubbles that
were formed along the electrode on the pressure side of the
generator were swept around the tip to form a relatively large
vortex core. The bubbles that were produced along the free-
stream electrode near the generator tip are seen to form the
outer helical structure of the vortex. Since the core of the
vortex leaves the trailing edge of the generator at a slightly
inboard location, the central portion of the vortex passes
above the airfoil even when the airfoil is at a small negative
incidence. Furthermore, it appears that the vortex survives its
encounter with the airfoil over an incidence range from
about -2° to +8°. At +9° incidence, however, the buffeting
effects of the separated flow over the trailing edge of the air-
foil causes the vortex to become unstable. At +10° incidence
the flow separates from the leading edge and causes the vor-
tex to become unstable before reaching the trailing edge of
the airfoil. This instability appears to grow until the vortex
becomes unrecognizable after passing about one airfoil-chord
length into the wake. As the airfoil incidence increases
further, the distance over which the vortex can still be recog-
nized behind the trailing edge of the airfoil decreases.
Because of the irregular and large-scale structure of the wake
behind the airfoil during static stall conditions, the interac-
tion between the vortex and the airfoil should be considered
an unsteady process.

The streamlines of the flow ahead of the airfoil are also
affected by the presence of the vortex. However, consider-
able care • must be taken when interpreting these results
because the vortex imparts a helical component to the flow
field, as a result of which the streamline visualizations
nowhere represent a two-dimensional cross section of the
flow. Accordingly, these results must be interpreted with
caution. Considering the airfoil at an incidence of +8°, and
comparing the weak -vortex case (fig. 8) with the no-vortex
case (fig. 7), it is apparent that two major changes have taken
place. First, the vortex (which is rotating counterclockwise

when viewed along a downstream direction) has lifted the
neighboring flow ahead of the airfoil (on the upwash side of
the vortex) by one streamline; and second, the separated
zone over the rear portion of the airfoil has increased greatly.
Comparing this flow with that for the case without a vortex
(fig. 7), and focusing on the airfoil at +10° incidence, sug-
gests that the effect of the vortex is to induce an increase in
the angle of attack by approximately +2° (based on the
amount of separation present in each case). Recalling that
these observations are applicable only to the upwash side of
the helical flow, it is important to note that a similar (though
not visible) but opposite condition must be occurring on the
downwash side. Since the core of the vortex not only appears
as a dense band of bubbles, but is central to the vortical
motion, an evaluation of its trajectory is more straightfor-
ward. The vortex core seems to move inboard from the
generator tip as it approaches the airfoil, cutting across the
streamlines that occur in the no-vortex case (see fig. 7 for
-8° and +8° incidence); but after reaching the suction peak
on the airfoil, the core closely follows the no-vortex stream-
lines. At an incidence of -2° (fig. 8), the outer part of the
vortex interacts strongly with the flow along the pressure
side of the airfoil. The vortex core is still visible, but the
outer helical streamlines disappear and instead become a
cloud of bubbles. At more negative angles of incidence, the
vortex becomes even more disorganized as it is pulled toward
the airfoil. When the airfoil is at -8° incidence, the vortex
nearly impacts on the pressure side of the airfoil close to the
leading edge. However, for more negative angles of incidence,
the vortex is driven away slightly from the airfoil surface. In
addition, an instability of the vortex core progresses
upstream from the wake (at -10° incidence), to the trailing
edge (-1 1°), and finally to a point ahead of the airfoil (-12°).
Rotating the generator to 10° incidence causes a much
stronger vortex to be produced (fig. 9). Although the trend is
essentially the same as that observed for the weak -vortex
case, certain features can be described with greater clarity
because of the more conspicuous behavior of the flow. In
comparing the weak-vortex flow field (fig. 8) with that
occurring for the strong vortex (fig. 9) when the airfoil is at
zero incidence, several observations can be readily made.
First, the bubbles comprising the vortex core are confined to
a more slender filament, no doubt a result of a greatly
reduced static pressure along the vortex core. Second, and in
keeping with a vortex of greater strength, the streamlines
that form the outer helical portion of the vortex are clearly
twisting at a much higher angular rate. Third, the core of the
vortex continues to leave the generator at about the same
slightly inboard position (0.09 c above centerline-grid line),
in spite of the difference in vortex strength. With regard to
the stability of the vortex core over the positive incidence
range of the airfoil, there is no significant difference between
the weak and strong vortex cases. The main difference
between the two cases occurs in the streamlines ahead of the
airfoil. Referring to the +8° of incidence case, for example,

the strong vortex (fig. 9) causes the neighboring flow ahead
of the airfoil (on the upwash side of the vortex) to be lifted
by two streamlines (compared to the no-vortex case, fig. 7),
whereas the weak vortex (fig. 8) shifted the flow by only one
streamline. In terms of induced separation over the airfoil,
the sequence of flows shown in figures 7-9 indicates that
separation occurs at slightly over 9° in the presence of a
strong vortex, at slightly under 10° for a weak vortex, and
probably at about 1 1° when no vortex is present.

With regard to the trajectory of the core of the vortex in
the +8 of incidence case, for example, there appears to be
no difference between the weak- and strong-vortex cases.
Although core instabilities were observed in the weak -vortex
case for -4° of airfoil incidence, their appearance is even
more striking during the strong-vortex interactions. Whereas
the core never quite collided with the airfoil in the weak-
vortex case, a direct impingement occurs at -6° of incidence
in the strong-vortex case. Direct impingement causes a wide
band of bubbles, with no apparent organized structure, to
appear in the wake of the airfoil. Continuing to focus on the
strong-vortex case, at -8° of incidence some degree of
periodicity can be seen in, the wake flow after passing over
the suction side of the airfoil, and, at -10°, the scale of this
periodicity increases. At -11° of incidence, a particularly
interesting event occurs. The core of the vortex just ahead of
the airfoil appears to undergo a helical distortion that is char-
acteristic of an unstable vortex. After colliding with the air-
foil, the flow breaks down over the pressure side of the
airfoil and is shed into the wake with a clearly periodic
organization (about 11.5 Hz). At -12° of incidence, the loca-
tion of this presumed vortex instability moves upstream
about one half of a generator-chord length ahead of the air-
foil. A similar breakdown of the vortex has been reported in
a smoke visualization test (ref. 8) of a vortex impinging on a
downstream airfoil.

Flow Visualization at Re = 120,000

The tip of the vortex generator was located on the cen-
terline of the tunnel, as well as offset from the centerline a
distance equal to one half of the generator chord. The vortex
generator was set to three angles of incidence, a = 0 , 5 ,
and 10°; and for each of these angles the downstream airfoil
was again varied from -16° to +16° (figs. 10-12). The
increase in Reynolds number for these results was obtained
by doubling the free-stream velocity. Since the duration of
the light pulse was not changed during this test, the particle-
path lengths at this higher speed will appear twice as long.
Although this streaking effect tends to lessen the clarity of
the in viscid portion of the flow field, it will aid in the identi-
fication of turbulent and rotational motions that occur in the
viscous portion of the flow field. In addition to the short-
duration exposures obtained using the strobe, long-time
exposures (20 sec) of the flow were made using a continuous

light source. The main purpose of the long exposures was to
obtain an accumulated visual record of the trajectory of the
vortex core in order to distinguish between regions having a
concentrated and well-defined vortex path and those where
lateral excursions and possible instabilities are present.
Included in these visualizations is a section of the boundary
layer exposed by hydrogen bubbles that were generated
along the chord of the airfoil. These bubbles are believed to
have had no measurable effect on the interaction. Although
the evidence is not conclusive, it was observed that as the
number of bubbles was increased (by increasing the voltage
on the electrodes) for photographic purposes, no change was
observed in either the thickness of the boundary layer or in
the proximity of the vortex to the airfoil. However, the
meaning of these visualizations requires some consideration.
Since the region of the interaction between the vortex and
boundary layer is known to be highly three dimensional, the
fact that only a narrow spanwise portion of the boundary
layer was visualized should be kept in mind when interpret-
ing the results. In the following discussions, the short-
exposure results will be addressed first, and in more detail.

Vortex generator on centerline- Once again, by placing
the generator at 0° incidence, a control case (fig. 10) was
produced against which comparisons could be made. There
are essentially no differences between these results and those
obtained at the lower Reynolds number, except that a more
definite Karman-vortex street can be detected in the bubbles
emanating from the trailing edge of the generator.

Rotating the generator to 5° incidence produces a vortex
core that is more visible (fig. 11) than the one obtained
under the same conditions at the lower Reynolds number
(fig. 8). The presence of a more visible core could be caused
by either a vortex of greater strength (therefore attracting
more bubbles because of the lower static pressure) or a visual
reinforcement of the core filament because of the streaking
allowed by the finite-time exposure. Another distinction is
that the neighboring flow ahead of the airfoil is shifted
upward by about one additional streamline (compare, for
example, the +8° of incidence flows in figs. 8 and 11). This
additional uplifting of the streamlines could be a result of
either a vortex of greater local strength (to be discussed
momentarily) or an increase in the size of the vortex so that
the region of high rotational velocity has moved farther away
from the center. In all other respects, however, the trends
observed earlier at the lower Reynolds number with regard to
the stability of the vortex core and the induced separation
over the airfoil remain essentially the same. One interesting
behavior that appears to be more distinct at the higher
Reynolds number concerns the vortex instability over the
airfoil. When the airfoil stalls at +10° of incidence (fig. 11),
the vortex core appears to undergo a more obvious helical
twisting motion.

The explanations given above for the additional uplifting
of the streamlines when the free-stream velocity was

increased may require some further discussion. Based on
classic aerodynamic theory, the swirl angle of a fully devel-
oped vortex can be argued to be independent of the free-
stream velocity.To demonstrate this point, consider the expres-
sions for the circulation on the generator, F - C^Uodc/2,
and the circumferential velocity component of an inviscid
vortex, w = r/4nr. These two equations can be combined to
obtain w = C^U^c/Sirr. An approximation for the swirl angle
can, therefore, be given by 6 « w/U^ = Cyc/8-nr, which is
independent of the free-stream velocity. To some extent this
conclusion is inexact because of the neglect of viscous
effects. It is more significant, however, that the arguments
given cannot be strictly applied in the vortex-development
region behind the generator. The extent of this development
region for a rectangular planform has been shown to be
about 4 chord lengths behind the generator (ref. 9). During
that time, reported measurements of the maximum circum-
ferential velocity of the vortex indicated that the swirl angle
decayed approximately 50% before the roll-up was complete.

Rotating the generator to 10° incidence produces a flow
(fig. 12) that, except for the differences noted above for the
5° case, is quite similar to that observed at the lower
Reynolds number (fig. 9). The maximum theoretical circula-
tion on the generator for this case is V = 0.28. Some of the
events that are more obvious in the higher Reynolds number
visualizations (fig. 12) concern the vortex impingement at
angles below -9° incidence. The region of vortex instability
ahead of the airfoil from -9° through -16° incidence is much
more pronounced. In addition, the breakdown of the flow
on the pressure side of the airfoil into periodically shedding
structures (about 23 Hz) is even more evident.

The same range of conditions for the generator and air-
foil were considered for the long-exposure visualizations
(figs. 13-15). When the generator is placed at 0° of inci-
dence, the bubbles that were produced near the tip (side
opposite from view) are observed to leave the trailing edge
over a broad band (fig. 13), instead of in a straight line
directly downstream of the electrode. This band can also be
seen in the short-exposure results (although less distinctly)
and is due to the slight spanwise-pressure gradient that drives
the bubbles inboard over the generator surface and away
from the tip. It may be useful to note that this band of
bubbles provides a white background against which the black
trailing edge of the generator can be easily identified in the
photographs. Keeping in mind that these bubbles are all pro-
duced on the pressure side of the generator, the influence of
the vortex on the flow near the tip can be better appreciated
when it is realized that nearly all of these bubbles are drawn
around the tip and become a part of the vortex core on the
upper surface just as it leaves the trailing edge (fig. 14). This
sweeping of fluid around the tip is even more dramatic when
the strength of the vortex is increased (fig. 15). This increase
in swirl angle is probably caused by the upstream movement
of the origin of the vortex on the upper surface of the gener-
ator (ref. 10).

Aside from the helical trajectory of the path of the
strong-vortex core (fig. 15) that extends over a distance of
1 .5 c downstream of the generator, the path of the vortex
core appears to be well defined and two dimensional as long
as the viscous region around the airfoil is avoided (from 0° to
+8° incidence in figs. 14 and 15). Although the path of the
vortex seems to be two dimensional over this distance, it is
actually more likely that some amount of transverse move-
ment (normal to the plane of view) is present as the vortex
encounters the circulation field around the airfoil. In fact,
this type of transverse distortion of the path of the vortex is
clearly evident in the results obtained in a similar experiment
(ref. 11), which included a side view and a plan view of the
vortex-airfoil interaction.

At -4° incidence the vortex can be distinguished from
the boundary layer in the weak-vortex case (fig. 14); how-
ever, a large thickening (and probable weakening owing to
viscous effects) of the vortex appears to have resulted from
the interaction in the strong-vortex case (fig. 15). At more
negative values of incidence, some thickening (or meander-
ing) of the vortex can be observed upstream of the airfoil.
However, the condition (or even survival) of the vortex after
mixing with the highly dissipative flow around the airfoil is
not certain. Considering, on the other hand, positive angles
of incidence for which the airfoil stalls (at or above +12°),
interaction with the separated region clearly produces a wide
band of vortex trajectories above the airfoil. This band,
which appears to broaden as it moves downstream, is not to
be interpreted as a vortex "burst" similar to that occurring
over delta wings at high incidence. Short-exposure results
(discussed earlier) have already established this to be a region
in which the core usually still exists as a filament (although
not always a stable one). This band indicates the extent to
which the vortex is jostled during its encounter with an
inherently unsteady separated zone.

Vortex generator off centerline- By offsetting the tip
of the generator from the centerline of the tunnel a distance
of one-half the chord of the generator, a relatively mild
vortex-interaction environment is produced. Placing the
generator at 0° incidence (fig. 16), as before, provides a basis
of comparison with other cases. Considering that portion of
the flow where streamlines exist for both conditions (that is,
below the centerline of the tunnel), the flow appears to be
independent of the extent to which the generator and its
wake protrude into the test section. Although some distur-
bance to the flow moving around the generator tip is present
in both cases, it has no observable effect on the stall of the
downstream airfoil.

Rotating the generator to 5° incidence produces a vortex
that passes well above the airfoil for the entire incidence
range from -16° to +16° (fig. 17). Judging by the size of the
vortex core, as well as by the rotational rate of the stream-
lines near the core, the vortex corresponding to the off-
centerline case (fig. 17) may be weaker than the vortex in the

on-centcrline case (fig. 11). This reduction in tip-vortex
strength could be attributed to an increase in the shedding of
vorticity into the wake before reaching the tip (in short, an
aspect-ratio effect). Although the vortex in this case is rela-
tively weak and remote from the airfoil, it nevertheless
induces the flow ahead of the airfoil to be shifted upward by
about one-half streamline on the upwash side of the helical
flow (compare, for example, the streamlines at zero inci-
dence in fig. 16 with those in fig. 17). As long as the airfoil is
not stalled, the vortex-airfoil interaction has no effect on the
stability of the vortex core. Even when the airfoil stalls
(a > 10°) and the buffeting action of the separated wake
interacts strongly with the vortex, there is still no clear evi-
dence of an instability. Rather, the evidence seems to show
that as long as a strong shear layer is not encountered, the
vortex is able to withstand relatively large transverse pressure
gradients without becoming unstable (see, for example, -12°
incidence in fig. 17).

Rotating the generator to 10° incidence causes the
vortex core and its surrounding helical streamlines to become
more distinct (fig. 18). The flow ahead of the airfoil is now
shifted upward about one streamline on the upwash side of
the helical flow around the vortex core; this shift is about
twice that observed for the generator at 5° incidence. Again,
once the airfoil stalls, the path of the vortex core can be seen
to go through large undulations as it interacts with the sepa-
rated zone downstream of the airfoil. In some cases (note
+14° and +16°), the vortex core appears to experience an
instability.

The path of the vortex core during long exposures is
shown in figures 19-21. The characteristics of the vortex are
essentially the same as those in the close -encounter case with
regard to its persistence while moving through the pressure
field created by the airfoil. Since the viscous region around
the airfoil is completely avoided, the interaction of the
vortex with the airfoil is strictly potential. Once again, when
the airfoil stalls, the boundary of the separated zone is
unsteady and causes the core of the vortex to be buffeted
over a band of trajectories.

Load Measurements at Re = 120,000

Lift, drag, and pitching-moment loads were measured at
a Reynolds number of 120,000. Data were taken at 1° incre-
ments of airfoil incidence over a range from -16° to +16°.
Because of the high density of data points, symbols have
been omitted from many of the figures in order to allow a
better examination of the curves that were constructed using
straight-line connections between the points.

Of initial concern was the unavoidable presence of the
generator wake and its possible effect on the loads of the
downstream airfoil. Although the greatest disturbance to the
flow field by the trailing-edge wake is created when the
generator is placed at maximum incidence (a = 10°), its

influence on the airfoil loads cannot be separated from the
more dominant effects of the tip vortex. The generator was,
therefore, placed at zero incidence in order to produce a
wake (albeit small), as well as a distortion of the flow around
the tip (but without producing a vortex). The results, which
are presented in figure 22, show that the presence of the gen-
erator in the free stream has essentially no effect on the air-
foil loads, even when the generator extends to the centerline
of the tunnel. Since some level of disturbance can be
expected when the generator is at incidence, the orientation
of the generator in the flow field with respect to the down-
stream airfoil in the present experiment has the advantage of
placing the wake farther away from the airfoil than the
vortex.

Placing the generator at incidence can be seen to have a
definite effect on the airfoil loads, especially when the vortex
makes a close encounter with the airfoil (fig. 23). The vortex
causes the airfoil to experience an early stall and a reduced
(more narrow) drag bucket. Note that only the pitching
moment shows any significant change at angles below stall.
This is probably caused by the presence of a laminar separa-
tion bubble, which becomes distorted so as to cause only a
shift in the center of pressure. The behavior of this bubble,
which no doubt is responsible for the kink in the lift curve
and the non-zero slope in the moment curve over the
unstalled range, is thought also to cause the stall to be differ-
ent from what is observed at higher Reynolds numbers
(refs. 12-14). Although the proximity of the vortex to the
leading edge of the airfoil is quite dependent on the sense of
the airfoil incidence (figs. 1 1 and 12), the vortex passes over
the suction side of the airfoil at the point of stall and causes
the same degree of early stall for both positive and negative
values of incidence. Based on the onset of lift and moment
stall (which appear to be more distinct than drag stall), the
interaction causes an early stall by 1 .6° in the weak-vortex
case and by 2.3° in the strong-vortex case.

When the generator is off centerline, a more modest
encounter with the airfoil results (fig. 24). The effects of the
vortex interaction are greatly reduced over the unstalled
region, but the same trends are observed as in the strong-
interaction case (discussed above). Although there is a differ-
ence in the post-stall curves depending on whether the airfoil
is at positive or negative incidence, it is interesting that the
angle at which stall occurs does not appear to be affected by
which side of the airfoil (pressure or suction) the vortex is
on. The most significant difference probably appears in the
sense of the rolling moment; however, this quantity was not
measured in this experiment. In the present case the interac-
tion causes an early stall by 0.8° in the weak-vortex case and
by 1 .7 in the strong-vortex case.

Theory at Re = 120,000

In order to better represent the conditions of the experi-
ment, extra panels were added to the formulation to simulate
the presence of the upper and lower tunnel walls. All of the
computations were made for the close encounter, strong-
vortex case. In other words, the generator tip was considered
to be on centerline with an incidence of 10°. Comparisons
with the experiment were made at three angles of airfoil inci-
dence: a = +8°, +12°, and +16°. The calculated path of the
vortex core will be discussed first.

Considering the case for the airfoil at +8° incidence, the
computed results are shown in figure 25(a) in the form of
streamlines leaving the trailing edge of the generator and
passing over the downstream airfoil. The core of the vortex
(shown as a dashed line) was computed to be the centroid of
the circulation for the vortices in the tip roll-up. The encir-
cled points were obtained from the experiment by making
discrete-coordinate measurements along the mean trajectory
of the vortex core (from fig. 21). This comparison shows a
rather favorable agreement between theory and experiment.

The computation for the interaction with the airfoil at
+ 12° incidence is shown in figure 25(b). For this calculation,
wake-relaxation iterations were required to simulate the flow
separation from the leading edge. After three iterations, good
agreement with the experimental data was obtained ahead of
the airfoil. However, in passing over the airfoil, the agree-
ment remains good only when considering the inner bound-
ary of the band of possible trajectories (the upper and lower
boundaries are indicated by the two symbols at each loca-
tion). Nevertheless, the agreement is classified as being gener-
ally good over the entire encounter, since it is beyond the
scope of present-day codes to account for this type of
unsteady separation behavior. The region of greatest disagree-
ment is just downstream of the trailing edge of the airfoil,
where the theoretical core appears to be diverging from that
observed in the experiment. This may be attributable to the
fact that calculations of the details of the roll-up were termi-
nated before passing downstream of the airfoil.

Examining the results for the final case with the airfoil
at +16° incidence (fig. 25(c)), the comparison between
theory and experiment is not especially good. The calcula-
tions made with a "no-separation" restraint agree reasonably
well with the experimental results ahead of the airfoil; how-
ever, the agreement is poor in the region over the airfoil. A
second calculation, which allowed for separation on the air-
foil, shows a very different trend; however, the agreement
remains poor. Although some of the differences between the
theory and the experiment can be reduced by increasing the
panel density on the generator (ref. 15) as well as by
accounting for the initial vortex development over the sur-
face of the generator, it may be that the greatest improve-
ment will come from a better separation model for the flow
on the downstream airfoil.

Based on the VSAERO code, the computed lift, drag,
and pitching-moment coefficients for the three angles of air-
foil incidence are shown in figure 26. When the airfoil is
stalled, it is clear that the first-iteration calculation (which
assumes the flow is fully attached) is incorrect. However, the
second-iteration calculation (which allows for flow separa-
tion) is in much better agreement with the experiment at
+ 16°. With the airfoil at +12°, the code predicts a partial
span separation over the upper surface, whereas the flow was
apparently fully separated in the experiment. This difference
is probably a result of the level of free-stream turbulence in
the present experiment, as well as the strong buffeting char-
acter of the stall observed for this airfoil. A partial span sepa-
ration can occur under certain conditions, as was the case
reported in reference 1 1 .

5. CONCLUSIONS

1 . A vortex may survive distortions caused by modest
values of transverse and axial pressure gradients more easily
than it can shear along its axis.

2. Buffeting from a nearby separated region can ini-
tiate a vortex instability, with the path of the core itself
assuming a helical shape.

3. An encounter between the vortex and the airfoil
boundary layer causes the interacting flow to mix and
emerge into the wake with no apparent vortex structure.

4. When the vortex impinges along the stagnation
region of an airfoil (and becomes subject to a strong adverse
axial pressure gradient), the core of the vortex becomes
unstable ahead of the airfoil and is then transformed into a
segmented and periodic structure as it moves over the surface
of the airfoil.

5. The presence of the vortex was found to cause pre-
mature stall in every case in this experiment. The greater the
strength of the vortex and the closer the encounter, the
earlier the stall.

6. The extent to which early stall occurs appears to be
independent of whether the vortex is on the pressure or suc-
tion side of the airfoil.

7. The theoretical model considered in this study accu-
rately calculates the vortex trajectory and airfoil loads prior
to stall. After stall, calculations for the vortex trajectory do
not compare well with the experimental data; however, those
for the loads are acceptable.

REFERENCES

1. Smith, W. G.; and Lazzeroni, F. A.: Experimental and

Theoretical Study of a Rectangular Wing in a Vorti-
cal Wake at Low Speed. NASA TN D-339, 1960.

2. McMillan, 0. J.; Schwind, R. G.; Nielsen, J. N.; and

Dillenius, M. F. E: Rolling Moments in a Trailing
Vortex Flow Field. NASA CR-1 5 1961, 1977.

3. Cheeseman, I. C: Developments in Rotary Wing Air-

craft Aerodynamics. Vertica, vol. 6, no. 3, 1982,
pp. 181-202.

4. Ham, N. D.: Some Preliminary Results from an Investi-

gation of Blade-Vortex Interaction. AHS J., Apr.
1974.

5. Ham, N. D.: Some Conclusions from an Investigation of

Blade-Vortex Interaction. AHS J., Oct. 1975.

6. Maskew, B.; and Rao, B. M.: Calculation of Vortex

Flows on Complex Configurations. ICAS-82-6.2.3.,
13th Congress of the International Council of the
Aeronautical Sciences, 1982.

7. Maskew, B.: Prediction of Subsonic Aerodynamic Char-

acteristics — A Case for Low-Order Panel Methods.
AIAA Paper 81-0252, St. Louis, Mo., 1981.

8. Patel, M. H.; and Hancock, G. J.: Some Experimental

Results of the Effect of a Streamwise Vortex on a
Two-Dimensional Wing. Aeronaut. J., Apr. 1974.

9. Chigier, N. A.; and Corsiglia, V. R.: Tip Vortices-

Velocity Distributions. AHS 27th Annual National
V/STOL Forum, 1971.

10. Hoffman, J. D.; and Velkoff, H. R.: Vortex Flow over

Helicopter Rotor Tips. J. Aircraft, vol. 8, no. 9,
1971.

11. Mehta, R. D.; and Lim, T. T.: Flow Visualization of a

Vortex/Wing Interaction. NASA TM-

12. Nakamura, Y.; and Isogai, K.: Stalling Characteristics of

the NACA 0012 Section at Low Reynolds Numbers.
Technical Report of National Aerospace Labora-
tory, NALTR-1 75, 1969.

13. Nagamatsu, H.; and Cuche, D.: Low Reynolds Number

Aerodynamic Characteristics of Low Drag NACA
63-208 Airfoil. AIAA 13th Fluid and Plasma
Dynamics Conference, 1980.

14. Mueller, T. J.; and Jansen, B. J., Jr.: Aerodynamic Mea-

surements at Low Reynolds Numbers. AIAA 12th
Aerodynamic Testing Conference, 1982.

15. Maskew, B.: Predicting Aerodynamic Characteristics of

Vortical Flows on Three-Dimensional Configura-
tions Using a Surface-Singularity Panel Method.
AGARDCP-342,1983.

c
c

t-l

c
a
.3
o

3
00

cc
O

\ "P If

§5

o
-a

a
o
•a

3
C

—

3
00

Figure 3.— Orientation of vortex generator and downstream airfoil in test section.

I-

-I
LU

z
g

i-
o
<

I-

cc
O

(0
O

-I
LU
CC

< LU

O cc

I-
o

III

f/1

c/3^

LL.

o —

G
00

3
00

z
o

-J

£
o

c
2

^3
00

13
C
r3

*o
13
O

-J
<

£>
■§

•O

C
<D
00
O

-a
>>

Figure 7.— Visualization of flow at Re = 60,000 with generator on centerline and set at a - 0° (no-vortex case).

Figure 7 — Continued.

Figure 7.- Continued.

Figure 7.— Continued.

5 = 0°

~-r — -•- -> r-

a =-11°

-'■

___

a = o°

_| L 1 .

—

~- • — — -T-. : --t- ■ '~ -+r '■ ■** ■ :"

a = -16°j

! StS*

•

Figure 7.— Concluded.

|Jr » Mpg g*" ■'■» £gS

Figure 8.— Visualization of flow at i?e = 60,000 with generator on centerline and set at a = 5° (weak-vortex case).

■Ml ■mm"' °Si imn ^•^■Mfc'

Figure 8.- Continued.

gSgH^Su.

PPJBB55

puni

Figure 8.— Continued.

a = 5°

a= -11'

:ss^S

■■■- l ■- — =]

■

a=5°

a= -16°

iiiy-'

. ■ ■ &■%

—^

*j^.-. ..-J

' jp

i S

llli

:
—

•

Figure 8.— Concluded.

Figure 9.- Visualization of flow at Re = 60,000 with generator on centerline and set at a - 10° (strong-vortex case).

Figure 9.— Continued.

a= 10°

— » — i

a = 0°

■ 1 — i —

B !■■■■■■■■■■■
•^^^^^^"^^^^^ ' i i "■ , - ^>W its ii

Figure 9.- Continued.

5= 10°

— i 1

—

a=-i0°

—! 1

fc„MK

Figure 9.— Continued.

k*| HELICAL DISTORTION

Figure 9.- Concluded.

Figure 10- Visualization of flow at Re = 120,000 with generator on centerline and set at a = 0° (no-vortex case).

rf «s

Figure 10.— Continued.

■

— ^^—

■

„

■

5=0°

^ ».

--. - .

w- ■

a = 16°

Figure 10 — Continued.

Figure 10.- Continued.

a = 0

■

Figure 10.— Concluded.

Figure 11.— Visualization of flow at Re - 120,000 with generator on centerline and set at a = 5° (weak-vortex case).

-#•5* - mi mmMm

Figure 1 1 .— Continued.

Figure 11.— Continued.

■— • - ,,- i>

■■■■■■■■■

=^5?

■■■■■■■■I
mmmmssmm

■BHmrf i ~

Bsauau
■■■■■■■■

Tt^r «^fe^.;i

Figure 11.— Continued.

■ggggjg^af! I

■■■IN in IhiMlliMMI

Figure 11.— Concluded.

5 = 10°

a = 10°

a = 2°

HUE

El ■■■i IE

WBIillUHl

iLililllllni

■■■■■■■■■■■■

Figure 12 — Visualization of flow at Re = 120,000 with generator on centerline and set at a = 10° (strong-vortex case).

«!

m§

K3iat>-I

g||H||UH|||aaai

Figure 12.— Continued.

■■■■■—«■

■■■■■■■■MP*

Figure 12.- Continued.

a= io°

— J^ 4

a= -4°

■Hi""

a= 10°

a=-8°

Figure 12.— Continued.

■

Figure 12 — Concluded.

^^ Hi HIH1 H1H1 HIH1 Hr7e9TLHI

Figure 13.- Long-exposure visualization of flow at Re = 120,000 with generator on centerline and set at a = 0° (no-vortex

case).

Figure 13 — Continued.

Figure 13.- Concluded.

II II II I^^H II W>^St

Figure 14.— Long-exposure visualization of flow at Re - 120,000 with generator on centerline and set at a - 5 (weak-vortex

case).

BB BB BB BB SI

Figure 14.— Continued.

- - - -

a = 5°

-r.

■■■Ml

mhhM

■p-

-20°

Figure 14.— Concluded.

Figure 15.— Long-exposure visualization of flow at Re = 120,000 with generator on centerline and set at a = 10° (strong-
vortex case).

Figure 15 — Continued.

a = 10°

■"■

_ 1

6--H

L^g

g*H

— 1

a =

■8°

^^^

■

mi mi

a = io°

Figure 15.— Concluded.

■

Figure 16- Visualization of flow at Re = 120,000 with generator off centerline and set at a = 0° (no-vortex case).

WBtm

Figure 16.— Continued.

—■■■—■

laif ■ si nun

!■■■■■■

Figure 16.— Continued.

Figure 16.— Concluded.

■■»

isir. ■■■!

HUHUHBHB

■■■■■■■■■■■■■■a

■BSHU

■■■■■■I

Figure 17.— Visualization of flow at Re = 120,000 with generator off centerline and set at a = 5° (weak-vortex case).

■!£>■■■■■■■■■■

Figure 17.— Continued.

• *****£?

— — v laaar

IknlBBBi

_-*#^*k ^-ii

■r

a=-10° ;

<S*^J<

Figure 17.- Continued.

mmmmmmm^^mmmmmmmmm

■■■■■■■■■■■■■■■■■a

Figure 17 — Concluded.

1
a =4°

■

__,. -

r. — — .._ -i

wmm

Figure 18.- Visualization of flow at Re = 120,000 with generator off centerline and set at a = 10° (strong-vortex case).

wbMM

Figure 18 — Continued.

UNSTABLE L J

Figure 18.— Continued.

is is il il ■■ II

■■■■■■■■■■■

Figure 18 — Continued.

■

mpBMtum

mmamm

a = -16° i

1 — i~t

Figure 18 — Concluded.

L--.

—

E^^^H

'-'•'

1 1

| |

Imh

— __^

Figure 19.- Long-exposure visualization of flow at Re = 120,000 with generator off centerline and set at a = 0° (no-vortex

case).

i^——

_HEP

■1

■*

__«_—-

—

-^^

—

a= 16°

™

■EBB7

■ i_____fc..

_^1

a= 20°

^^^^

■

—

■

—

a= 0°

■■■■l""""^*

a = 0°

^^^^^gH

■_______■____-_

—

■

_________

■■■■_■

___i _____>-

■■

a= -4°

Figure 19.— Continued.

Figure 19 — Concluded.

a = 0°

Figure 20.— Long-exposure visualization of flow at Re = 120,000 with generator off centerline and set at a = 5° (weak-vortex

case).

5 = 5°

___

-—

—J

—^■^83

^^^^

- - - •

L—

^^■■B

^^■■■1

MM^JJ

- -- —

= 0°

Figure 20- Continued.

Figure 20.— Concluded.

r«\&w - - mil ■■ ■■ I ■■ ■■

Figure 21 .- Long-exposure visualization of flow at Re = 120,000 with generator off centerline and set at a = 10° (strong-
vortex case).

a= 10°

^■1

*-.

—

__.

ZT_

^^^

;

a=uu

Figure 21.— Continued.

•-

5= 10°

— f; -

"--

-----

— I — -

—I 1

fJam

^^^^"~

n o°

U —

-o

■■ ■■Ml

Figure 21 .— Concluded.

AIRFOIL ALONE

AIRFOIL AND GENERATOR

1.0 | 1 1 1 1 r

L 0 -

CD-1

-.1

.10

.05 -

CM °

^\\ ' - " ^0*

\ ■ f

V /

-.10

-15 -10 -5 0 5

(a) Generator on centerline.

- •

(a.

X I

: \*r

<u ;

10 15 -15 -10 -5 0

oi, deg

(b) Generator off centerline.

5 10 15

Figure 22.— Generator wake effects on airfoil loads when a - 0 .

GENERATOR FEATHERED

GENERATOR AT INCIDENCE

L 0 -

CD .1 -

o -

•M o -

-.05 -

REDUCED DRAG
BUCKET

CHANGE IN BUBBLE

EARLY MOMENT
STALL

(b)

15 -15 -10

oc, deg

-5

10 15

(b)a=10°.

Figure 23.— Airfoil loads during vortex encounter with generator on centerline.

CL 0 -

.2 ^

0 -

-.1
.10

0 -

-.05 "

-.10

GENERATOR FEATHERED
GENERATOR AT INCIDENCE

V 4

_ \t : j/. . -

v '/

X. ' J

■ v. Y

-15 -10 -5

(a) a = 5<

11 ;

1 v

1 t

1 * --***

1 \ ^^!^>

\*r

S*^~ 1- 1

1 1
It 1

■ * 1

(a) :

1 1 1 1 1

EARLY LIFT STALL

n 1 1 1 r

CHANGE IN BUBBLE

EARLY MOMENT
STALL

10 15 -15 -10

a, deg

-5

\ \
_ V

I I I

REDUCED DRAG

BUCKET

i 1
tjr

-* ^~^*

I I I I I

10 15

(b)a=10°

Figure 24.- Airfoil loads during vortex encounter with generator off centerline.

POINTS MEASURED FROM VORTEX TRACK
IN EXPERIMENT

NO SEPARATION \ COMPUTED CENTROID

WITH SEPARATION MODEL 0F VORTICITY LOCUS

(a) a = +8°.

(b)a = +12c

(c)a = +16°.

Figure 25- Comparison of theory and experiment.

CM 0 -

VSAERO
CALCULATION

Figure 26.— Comparison of theory and experiment over separated region.

1 Report No.

NASA TP-227 3 AVSCOM TR 83-A-17

2. Government Accession No

3 Recipient's Catalog No.

4 Title and Subtitle

AIRFOIL INTERACTION WITH AN IMPINGING VORTEX

5. Report Date

February 1984

6. Performing Organization Code

7 Author(s)

K. W. McAlister and C. Tung

8. Performing Organization Report No.

A- 954 3

9. Performing Organization Name and Address

Aeromechanics Laboratory

USAAVSCOM Research and Technology Laboratories

NASA Ames Research Center

Moffett Field, CA 94035

10. Work Unit No.

K-1585

11. Contract or Grant No.

12 Sponsoring Agency Name and Address

National Aeronautics and Space Administration
Washington, DC 20546

and
U.S. Army Aviation Systems Command
St. Louis, MO 63166

13. Type of Report and Period Covered

Technical Paper

14. Sponsoring Agency Code

992-21-01

15 Supplementary Notes

K. W. McAlister and C. Tung: Aeromechanics Laboratory, USAAVSCOM Research
and Technology Laboratories.

Point of Contact: K. W. McAlister, Ames Research Center, MS 215-1, Moffett
Field, Calif. 94035 (415)965-5892 or FTS 448-5892

16. Abstract

The tip of a finite-span airfoil was used to generate a streamwise
vortical flow, the strength of which could be varied by changing the inci-
dence of the airfoil. The vortex that was generated traveled downstream
and interacted with a second airfoil on which measurements of lift, drag,
and pitching moment were made. The flow field, including the vortex core,
was visualized in order to study the structural alterations to the vortex
resulting from various levels of encounter with the downstream airfoil.
These observations were also used to evaluate the accuracy of a theoretical
model.

17. Key Words (Suggested by Author(s) )

Vortex interaction
Vortex instability
Airfoil stall

18. Distribution Statement

Unclassified

Unlimited

Subject Category: 02

19. Security Oassif. (of this report!

Unclassified

20. Security Classif. (of this page)

Unclassified

21. No. of Pages

22. Price"

A02

'For sale by the National Technical Information Service, Springfield, Virginia 22161

NASA-Langley, 1984

National Aeronautics and
Space Administration

Washington, D.C.
20546

Official Business

Penalty for Private Use, $300

THIRD-CLASS BULK RATE

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National Aeronautics and
Space Administration
NASA-451

l\J/\SA

If Undeliverable (Section 1S8
Postal Manual) Do Not Return

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