An experimental investigation was carried out to study the aerodynamics of an airfoil with a rectangular jet exiting from its lower surface at fifty percent of the chord. The airfoil was tested with and without the influence of a ground plane. Surface static pressures were measured on the airfoil at jet to free stream velocity ratios ranging from 0 to 9. From these pressures, the variation of C sub L with velocity ratio was easily determined. The measurements indicated significant positive and negative pressure regions on the lower surface of the airfoil ahead of and after the nozzle exit respectively. The presence of a ground plane enhanced these pressure regions at low velocity ratios, but at a particular ratio for each plane location, a recirculation zone or vortex formed ahead of the jet resulting in decreased pressures and a drop in C sub L

Krothapalli, A.

Leopold, D.

English

NASA Technical Reports Server

N89-10855
EFFECTS OF A GROUND VORTEX ON THE AERODYNAMICS OF AN AIRFOIL
by
A. Krothapalli
Fluid Mechanics Research Laboratory
The Florida State University
Tallahassee, FL
and
D. Leopold
Department of Aeronautics & Astronautics
Stanford University, Stanford, CA
ABSTRACT
An experimental investigation was carried
aerodynamics of an airfoil with a rectangular
its lower surface at fifty percent of the chord.
out to study the
jet exiting from
The airfoil was
tested with and without the influence of a ground plane.
static pressures were measured on the airfoil
stream velocity ratios
pressures, the variation
determined.
Surface
at jet to free
ranging from 0 to 9. From these
of C L with velocity ratio was easily
The measurements indicated significant positive and negative
pressure regions on the lower surface of the airfoil ahead of and
after the nozzle exit respectively. The presence of a ground
plane enhanced these pressure regions at low velocity ratios but
at a particular ratio for each plate location, a recirculation
zone or a vortex formed ahead of the jet resulting in decreased
pressures and a drop in C L.
147
https://ntrs.nasa.gov/search.jsp?R=19890001484 2020-03-20T06:06:39+00:00Z
During
interest
jets and
fuselage.
INTRODUCTION
the past several years there has been increased
in V/STOL aircraft configurations which utilize lift
thrust augmentors mounted in the wings and/or the
One such configuration of interest uses a high lift
system consisting of a wing with a long rectangular jet along the
span issuing from below. Such a jet could be produced by
installing two dimensional ejectors along the span of the wing.
While these configurations usually exhibit improved lift
characteristics, the interaction of the jet and the free stream
can result in undesirable aerodynamic loading characteristics
influencing the aircraft performance. For example, in hovering
entrainment of the surrounding air by the jet induces a suction
pressure on the lower surface of the wing causing a downward or
suck-down force. During the transition from hovering to
conventional forward flight this interaction produces a region of
positive pressures upstream of the jet and a region of negative
pressures downstream of the jet resulting, under certain
conditions, in a net loss of lift and a nose-up moment. When the
aircraft is operating in STOL mode, all the induced effects
discussed above are present but modified by the presence of the
ground. Close to the ground, the jet impinges on the ground and
forms a wall jet that flows outward from the impingement region.
The wall jet formed upstream of the jet exit, rolls up forming
what is commonly known as a "ground vortex". This is a result of
the interaction of the wall jet with the oncoming free stream.
The location of this ground vortex and its induced effects on the
148
nearby
performance of the STOL aircraft.
Various aspects of the jet induced
fuselages have been the subject of many
lifting surfaces is of importance in predicting the
effects on wings and
studies; and most of
these have been experimental investigations. Currently, in most
V/STOL aircraft designs, a semi-empirical approach guided by
experimental data is followed to model the specific jet-induced
flow field. Several researchers, over the years, have surveyed
and described these jet-induced or propulsive effects (Margason I ,
Skifstad2). More recently Kuhn 3 gave a comprehensive account of
the induced aerodynamics of jet and fan powered aircraft. And
recent advances in prediction methods for these effects on V/STOL
aircraft were described by Agarwal 4 . Since these reviews are
quite extensive, no attempt is made here to discuss the previous
work on jet induced aerodynamics.
The problem addressed here is the determination of the
various aerodynamic forces of the airfoil resulting from a jet
issuing normal to its chord line into a uniform cross flow and in
the presence of a ground plane, as shown in Figure i. The
interaction between the jet and the cross flow in the presence of
an airfoil is characterized by the following parameters: the
geometry of the airfoil, angle of attack of the airfoil, free
stream Mach number, free
chord of the airfoil, the
location and orientation
stream Reynolds number based on the
geometric parameters of the nozzle,
of the nozzle with respect to the
airfoil, Mach number of the jet, the location of the ground plane
with respect to the airfoil and the nature of the conditions at
149
the nozzle exit and the free stream.
The airfoil used was the NACA0018. A rectangular nozzle of
aspect ratio 87 was selected for the nozzle and the nozzle was
oriented lengthwise along the span. The nozzle was located at 50
percent of the chord. The exit section of the nozzle was
designed such that the jet exits normal to the chord. The free
stream velocity was varied from 20m/sec to 60m/sec. The
corresponding Reynolds number R - U C/_ , varied from 2 x 10s to
6 x 10 s . The mean velocity at the nozzle exit was varied from 20
m/sec to 250m/sec. The angle of attack of the airfoil was kept
at zero degrees.
APPARATUS, INSTRUMENTATION AND PROCEDURES
The wind tunnel used in this experiment was subsonic closed
circuit type. The test section has the dimensions of 90.2 x 45.7
x 45.7cm. The flow speed in the test section can be varied
between 20m/sec and 65m/sec. The model was situated midway
between the upper and lower walls of the test section.
A NACA 0018 symmetric airfoil was chosen for the experiment.
The airfoil was made in several sections using aluminum and
stainless steel. It has a 15cm chord and spans the entire 45.7cm
width of the test section. The aspect ratio of the wing was
therefore equal to about 3.05. A rectangular slot with its long
dimension in spanwise direction was cut into the lower surface at
midchord. The length and width of the nozzle exit were 25cm and
0.3cm respectively. Before air reaches the nozzle exit, it
passes through a settling chamber placed inside the wing and
150
extends along the span of the wing. Compressed air was supplied
to the settling chamber from both ends of the wing. To ensure a
uniform flow at the nozzle exit, adjustable vanes were placed
inside the settling chamber. The inlet section of the slot was
designed in such a manner that the jet stream exhaust
perpendicular to the chord of the airfoil. With the optimum
position of the vanes, a uniform flow was obtained. The
variation of the mean velocity along the span was within ten
percent of the value at the center of the nozzle exit. The jet
exit velocity was varied from 0m/sec to 240m/sec. For simulation
of a ground plane, an aluminum plate of 45.7cm wide, 61cm long
and 0.3cm thick was used. The leading edge of the plate was
rounded into parabolic shape and a flap was attached at the
trailing edge to insure a attached flow at the leading edge. A
transition strip was placed 10cm from the leading edge. The
distance between the ground plane and the airfoil chord was
varied from 3.75cm to 15cm.
Surface pressure measurements were made at several velocity
ratios (nozzle exit mean velocity/free stream mean velocity). To
obtain the aerodynamic force coefficients, the surface pressure
data was integrated around the airfoil at mid-span location. The
jet reaction is not included in most of the datalift due to the
presented.
A Cartesian coordinate system (x,y,z) was used with the x-
axis oriented along the center line of the wing section and with
the origin located at the leading edge as shown in figure i.
For most of the measurements errors were estimated to be of
151
the order of five percent.
The testing of V/STOL models in wind
problems that are not encountered
conventional airfoils, where the
relatively well understood 5 • V/STOL
tested have a relatively large wake
tunnel presents many
in the testing of the
testing techniques are
models such as the one
deflection angle which
presents one of the most difficult problems that is encountered
in wind tunnel testing. The primary work on wind tunnel wall
effects and their corrections for V/STOL configurations was done
by Heyson 6 . Studies covering the limits on the minimum speed in
V/STOL wind tunnel test were done by Raev . Recently Margason and
Hoad6 gave an account of V/STOL aircraft model wind tunnel
testing from model design to data reduction. In most of the
instances, the model used is a fan-in-wing configuration. Since
these correction techniques are highly configuration dependent,
and the present wing model is not representative of any flight
vehicle, no attempt is made here to correct the data.
One particularly important aspect of V/STOL model testing is
the need to describe a "jet-off reference configuration" for each
jet-on configuration tested.
a basis for determining
aerodynamic characteristics.
experiment.
Another factor to
on the ground plane.
These data are then used to provide
the interference of the jets on
Such a procedure was used in this
take into account is the flow impingement
In a wind tunnel with the air moving with
respect to the model and to the ground plane, there is a boundary
layer on the floor. The effect of this can be minimized by using
152
moving belt ground plane. Several investigationsa
carried out on this subject by Hackett et al 9 .
experiments described here the ground plane was fixed
attempts have been made to bleed the boundary layer.
have been
In the
and no
RESULTS AND DISCUSSION
Typical variations of the surface pressure on both upper and
lower surfaces of the free airfoil, at zero angle of attack with
and without the jet, are shown in Figure 2. The pressures are
plotted in the form of the pressure coefficient Cp given by
Cp _ (p - p.)/q.
It is observed that without the jet,
distribution on both sides of the airfoil are
identical, confirming the symmetric property of the airfoil. For
a velocity ratio (jet exit velocity/freestream velocity) of 6,
the pressure
very nearly
the influence of the jet on the surface pressure is quite
significant as shown in the figure. When comparing this
distribution with the jet-off condition, the following
observations are made: on the lower surface, in the region
upstream of the jet, an increase in pressure occurs, while a
decrease in pressure is noticed in the region behind the jet.
jet is a result of the
The effect of this is an
of the airfoil, resulting
The positive pressure ahead of the
blockage of the free stream by the jet.
increase in effective angle of attack
in a relatively low pressure on the upper surface of the airfoil.
At very low velocity ratios, the recirculation zone behind the
jet is small and the flow reattaches to the lower surface. As
153
the jet strength is increased by increasing the exit velocity,
the flow in the region between the jet and the trailing edge
forms a recirculation region and it extends into the wake. The
magnitude of the pressure coefficient in this region was observed
to be fairly constant as depicted by its distribution in the
region between the jet and the trailing edge in figure 2. The
"Kutta Condition" requires that the pressure on both lower and
upper surfaces at the trailing edge be equal. This being the
case, the pressure on the upper surface near the trailing edge is
fixed by its value in the recirculation region on the lower
surface or vice-versa., It is interesting to note that very
little variation in the magnitude of Cp is observed on the upper
surface for x/c greater than about 0.6, thus suggesting that only
the pressure changes in the first half (x/c < 0.5) of the airfoil
are mostly responsible for the generation of the induced lift.
From these observations it may be suggested that the positive and
negative pressure regions on the lower surface are essentially
responsible for many changes in the gross aerodynamic
characteristics of the airfoil to be noted later.
the chordwise pressure distribution determined, theFrom
sectional
pressure
velocity
lift was easily obtained by
over the span wise section.
ratio in the range tested, a
numerical integration of
For each value of the
corresponding sectional
lift
reaction force
of the lift.
velocity ratio.
coefficient C L was obtained at mid span location. The
due to the jet is not included in the definition
Figure 3 shows the variation of C L with the
It can be seen from the figure that the C L
154
increases monotonically up to a velocity ratio of about 5. This
is a result of the pressure increasing rapidly in front of the
jet on the lower surface. In the range of velocity ratios
between 5 and 8, the pressure in the recirculation regions behind
the jet decreases rapidly with increasing velocity ratio, thus
resulting in a drop of C L as shown in the figure. The detailed
discussion of these regions and their effect on the aerodynamics
of a free airfoil is given by Krothapalli and Leopold I° . For
velocity ratios greater than 8, an increase in C L is observed,
and this is attributed to the influence of the tunnel wall or
ground effect.
At low velocity ratios, the influence of a ground plane on
the flow around the airfoil seems
trends found for the free airfoil.
behind the jet both increase and
to enhance the same general
The regions ahead of and
decrease respectively with
increasing velocity ratio but the variations are more pronounced;
the degree of which depends strongly on plate position. This
phenomenon is shown in figure 4 where the pressure distribution
for two plate locations are compared to the free airfoil
distribution at a velocity ratio of 2. As this ratio is
increased, the pressure in front of the jet drops dramatically
resulting in a sharp decrease in C L. The velocity ratio at which
this occurs depends strongly on plate position.
The variation of lift coefficient throughout the velocity
ratio range is shown in figure 5 for three plate locations and
are compared with the free airfoil. These curves indicate a
unique velocity ratio for each plate location where C L reaches a
155
maximum. As the ratio increases further, the pressure in front
of the jet on the lower surface decreases thus resulting in a
reduced CL. These unique points represent a boundary between
favorable and unfavorable operating conditions which are shown in
figure 6. In region i, favorable conditions exist.
coefficient increases as the jet velocity
II, unfavorable conditions exist since
decreases as the jet velocity increases.
6 also indicates the beginning of a new type of flow structure in
front
aerodynamics
increases.
where case
of the jet; the influence of
of the airfoil increases
This complex flow is best
I corresponds to the flow
The lift
increases. In region
the lift coefficient
The relation in figure
which on the overall
as the velocity ratio
visualized in figure 7
condition occurring in
region 1 of figure 6. For this case, the momentum of the jet is
small enough for the jet stream to be bent by the oncoming free
the region behind the
behind the jet and a
In case II, the jet
stream. The recirculation is confined to
jet therefore creating a low pressure zone
high pressure zone in front of the jet.
momentum reaches a high enough value that the jet impinges normal
to the plate creating two recirculation regions. The region in
front of the jet drops in pressure resulting in a decreased C c.
As the jet velocity increases, the recirculation zone in front of
the jet increases in intensity and eventually forces C L to a
negative value. The recirculatory region in front of the jet is
commonly known as the ground vortex.
156
CONCLUSIONS
From this preliminary experimental investigation, the
following conclusions can be drawn. The static pressure
distribution around the airfoil shows two distinct regions on the
lower surface, which greatly influence the overall aerodynamics.
First there is the positive pressure region upstream of the jet.
This is attributed to the "blockage" of the freestream by the
jet. The second is the region between the jet and the trailing
edge, marked by the negative pressure coefficient, and the
magnitude of the pressure coefficient in this region is found to
be nearly constant. The pressure on the upper surface of the
airfoil is also influenced by the presence of the jet, and the
influence is such that only the pressure distribution for the
leading half of the airfoil contributes to the lift coefficient.
The presence of the ground plane, for moderate heights, and at
low velocity ratios, improves the aerodynamic characteristics of
the airfoil. However, a further increase in the velocity ratio
for a fixed ground plane height, a large vortex develops in front
of the jet, commonly known as "ground vortex", resulting in a
sharp decrease in CL.
ACKNOWLEDGMENTS
The work presented here has been supported by NASA Ames
Research Center under grant numbers NAG2-111 and NAC2-198 and the
work was carried out at Stanford University.
157
REFERENCES
i. Margason, R.J., "Review of Propulsion Induced Effects on
Aerodynamics of Jet V/STOL Aircraft", NASA TN D - 5617, 1970.
2. Skifstad, J.G., "Aerodynamics of Jets Pertinent to V/STOL
Aircraft", Journal of Aircraft, Vol. 7, No. 3, 1970.
3. Kuhn, R.E., "The Induced Aerodynamics of Jet and Fan Powered
V/STOL Aircraft", Recent Advances in Aerodynamics, Edited by
A. Krothapalli and C.A. Smith, Springer-Verlog, 1986.
4. Agarwal, R.K., "Recent Advances in Prediction Methods for Jet
.
Induced Effects on
Aerodynamics, Edited
Springer-Verlog, 1986.
V/STOL
by A.
Pope, A. and Harper, J.J.,
John Wiley & Sons Inc., 1966.
i0. Krothapalli, A. and Leopold,
Airfoil with a Jet Issuing
Aircraft", Recent Advances in
Krothapalli and C.A. Smith,
"Low-Speed Wind Tunnel Testing",
6. Heyson, H.H., "The Effect of Wind Tunnel Wall Interference on
the Performance of a Fan-In-Wing VTOL mOdel", NASA TN D-7518,
1974.
7. Rae, W.H., "Limits on Minimum Speed V/STOL Wind Tunnel
Tests", Journal of Aircraft, Vol. 4, No. 3, 1967.
8. Margason, R.J. and Hoad, D.R., "V/STOL Aircraft Model in Wind
Tunnel Testing from Model Design to Data Reduction", Vol. 17,
No. 3, 1980.
9. Hackett, E.J. and Bles, R.A., "Ground Interaction and Tunnel
Blockage for a Jet-Flapped, Basic STOL Model Tested to a Very
High Lift Coefficient", NASA CR - 137, 1976.
D., "A Study of Flow Past an
from its Lower Surface", Joint
158
Institute of Aeronautics and Acoustics TR-54,
University, June 1984.
Stanford
159
¢NACA 0018
h
Figure i. Schematic of the model and nomenclature.
160
Cp
i _-'_-'-" - _ / 0 Upper Surface I
_'_// __ A Lower Surface
l(l_ vj/v_: o
vj/v_= 6
0 O.5 1
¥/_
Figure 2. Surface pressure distrib/ut_zon of the airfoil at midspan and
out of ground effect.
161
(I)
<_<_
F-,_I
0 1_-
i_.i-i
_-i-i_.
_< 0
0
i_.
t-h_
0
(I)I_.
0
0 1_-
t-hl-h
(-I-
(-I-
0
(I)
t-h_-h
_,-
I_._
I-,-I--,-
F,-
I I I
,%
l"--i ,--%
"_ r'D
CD i"_
_-.- ¢-.t
_0
_...._r--oe
I
!
/ d
\
\Q
Cp
Figure 4. Surface pressure distribution of the airfoil at midspan for
different ground plane locations; velocity ratio = 2.
1
163
2.00
1.75
! .50
Free airfoil
-----o------h/c = 0.75
h/c = 0.50
----.o--- h/c = 0.25
I I I I
1.25
1.00
C L 0.75
0.50
0.25
0
I I
2 4
I I
6 8
V R = (VjlVoo)
Figure 5. Variation of the jet induced sectional lift
coefficient with velocity ratio for different
ground plane locations.
0
164
Woo
10
8
6
4
2
0
0
(Flow
to
I I I
Region !!
structure corresponds
Case 11of Fig.l)
Region I
(Flow structure corresponds
to Case 1 of Fig.7)
I I I
0.25 0.50 0.75 I.O0
h/c
Figure 6. The loci of the maximum C L
plane locations.
for different ground
165
NACA 0018
h1.
_\\\\\\\\\\_,\\\\\\\\
Case !1
_\\\\\\\\\\\\\\\\\\\\
Figure 7. The flow structure corresponding to regions
identified in figure 6.
166


Effects of a ground vortex on the aerodynamics of an airfoil

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