using sweeping jets on swept wings - unina.it · tests were carried out at the university of...

15th International Conference on Fluid Control, Measurements and Visualization

27-30 May 2019, Naples, Italy

Paper ID: 294 1

Using Sweeping Jets on Swept Wings

Elisa Phillips1,*, Israel Wygnanski2

1Research Scientist, Department of Aerospace and Mechanical Engineering, University of Arizona, Tucson, AZ, USA 2Professor, Department of Aerospace and Mechanical Engineering, University of Arizona, Tucson, AZ, USA

*corresponding author: [email protected]

Abstract Tests were carried out on a 45° swept back wing that is highly unstable in pitch due to its large aspect

ratio and sweep-back angle. The tests were carried out at velocities of 20-40 meters/second at a representative free

stream Reynolds numbers approaching 106. The wing being based on a NACA0012 airfoil had a round leading

edge that prevented the generation of a leading edge vortex prior to tip trailing edge stall, thus its first pitch-up

departure was due to flow separation at the tip. The interaction between the tip separation and the leading edge

vortex resulted in another non-linear pitch behavior. Finally, at larger incidence angle the leading edge vortex

propagated inboard and the trailing edge separated region moved upstream. This wing would have been

uncontrollable in pitch, were it not for the introduction of active flow control. In this case a single sweeping jet

actuator could extend the trimmed incidence angle by 5° to 7°, thus avoiding the pitch break and almost doubling

the usable lift coefficient. Balance results, flow visualization and Pressure Sensitive Paint were used to correlate

the quantitative results to flow physics.

Keywords: Active Flow Control, swept back wing, leading edge vortex, pitch control

1 Nomenclature

AFC = Active Flow Control

AR = Aspect Ratio

b = wing span

CD = drag coefficient

CL = lift coefficient

CLL = rolling moment coefficient

CLM = pitching moment coefficient

CLN = yawing moment coefficient

c = chord

Cµ = dimensionless momentum coefficient

D = drag

L = lift

LE = leading edge

Mn = Mach number

MRL = Moment Reference Line

NP = Neutral Point

PSP = Pressure Sensitive Paint

Re = Reynolds number

S = wing reference area

TE = trailing edge

U∞ = free stream velocity

α = angle of incidence

β = jet axis relative to TE

Ʌc/4 = quarter chord sweep angle

Ʌ = leading edge sweep angle



Paper ID: 294 2

2 Background and Introduction

All subsonic jet aircraft have swept back wings in order to cruise efficiently at high Mach numbers. Only the

normal velocity component of the free stream is brought to stagnation at the Leading Edge (LE) due to the

sweep and it is this component that accelerates rapidly near the LE before decelerating again and potentially

stagnating near the Trailing Edge (TE). Therefore, the surface flow is dominated by the spanwise velocity

component near the LE attachment line and in the much broader TE region (Figure 1 right). Since deceleration

of fluid is synonymous with adverse pressure gradient the flow near the surface should turn outboard as it

selects the path of least resistance resulting in the accumulation of low momentum, vortical fluid near the wing

tip. Thus, a slight positive deflection of an aileron near the wing tip results in separation and renders such a

control surface ineffective (Figure 1 right). When separation occurs near the LE of a thin wing, it starts

somewhere near the tip but outside the zone of influence of the tip vortex and it is usually followed by

reattachment farther downstream. Such a local separation encloses a vortex (a bubble in two-dimensions) that

increases the lift generated by the wing. However, the origin of such vortex propagates inboard with increasing

incidence until it reaches the wing root or a discontinuity in the LE shape of the wing or its sweep. The inboard

propagation of the LE vortex and the increase in its strength with increasing incidence is accompanied by non-

linearity in the pitching moment and a potential loss of longitudinal stability. A small increase in incidence

angle may even change the sign of the pitching moment making the control of a potential airplane model very

difficult.

The research on swept-back wings peaked after World War II when jet propulsion enabled high-speed flight.

Scores of NACA research memoranda were published on a variety of wing planforms, thicknesses, sweep back

angles, Aspect Ratios (AR) etc. An extensive report tabulating the pitching moment characteristics of hundreds

of wing configurations is provided by Furlong & McHugh [1]. An example is provided in Figure 1 (left) where

the effect of sweep-back on pitching moment of a low aspect ratio wing, based on a NACA 65A006 airfoil, is

shown. It suggests that wings having larger sweep back angles, Λc/4, are difficult to trim and control over a

wide range of lift coefficients, CL. This is because many wing planforms are statically unstable whenever the

sweep or the aspect ratio or both are large. Most transport airplane wings designed during the past fifty years

have Ʌc/4≈30° and only their AR was allowed to increase in part due to the introduction of light & strong

composite materials.

Figure 1 Left: The effect of sweep on the pitching moment of a simple wing [1] at a given CL. AR=4, Taper ratio=0.6 Right: Oil flow visualization over a generic transport aircraft wing [2]



Paper ID: 294 3

Although many of today’s newest aircraft are controlled by computers (fly by wire designs), the aerodynamics

resulting in pitch instability at high lift is still an enigma requiring an innovative solution. Since a low

sweepback angle limits the cruise speed of an airplane some military airplanes capable of high speed

(supersonic dash) had swing wings (e.g. F-14, B-1 & MiG-23), but this solution is both complex and heavy.

Consequently, most current combat aircraft have low AR wings (e.g. F-22, F18 & F-35) with a variety of high

lift devices to improve maneuverability and prevent “wing drop”. They also have a very low Lift to Drag (L/D)

ratio and require aerial refueling in order to extend their range.

The pitch instability problems are more severe on tailless configurations because the control surfaces are never

far enough from the center of gravity, yet many such aircraft are currently being designed and tested because

of their aerodynamic efficiency and stealth qualities. A typical tailless configuration has a λ shape with a thin

outboard wing portion while its central body has to be thick in order to contain the propulsive and support

systems and a cargo bay as large as possible. Typical leading edge sweep-back varies between 45o-55º and a

prevailing AR varies between 3 and 5. The higher AR planforms perform reasonably well at cruise (L/D>10),

but they are problematic at high incidence angles needed during takeoff and landing. Furlong & McHugh [1]

report that nose-up pitch departure depends mostly on Λ and AR (or weighted average sweep measured at the

¼ chord, Λc/4, for tapered and cranked wing planforms). If the maximum lift coefficient, CLmax, is of order

unity, the maximum useable CL may be as low as ⅓ before reaching pitch-break, beyond which large control-

surface deflection may be required for trimming.

Figure 2: Longitudinal stability boundary as a function of AR & seep back of quarter chord



Paper ID: 294 4

The ESDU preliminary design data sheets [3] suggest that the product 2.6>AR*tan(Λc/4)>3.0 approximately

follows an empirical demarcation line separating stable from unstable wing configurations (Figure 2). Slats,

fences, vortilons, snags and LE cranks improve the pitch stability of a given wing planform but they add

complexity and weight to the model, as well as drag to its aerodynamic performance. One method of alleviating

or delaying pitch–break is to reduce AR, therefore increasing induced drag; another is to make the LE thicker

resulting in drag rise at high speed. A LE crank reduces Λc/4 while increasing AR and is thus a method of choice

on recent models, tested on Northrop-Grumman’s X-47C & Boeing’s X-48. Since most tailless aircraft

planforms are located on or near the neutral stability curve (Figure 2), pitch–break is a major impediment to

the design. Cruise performance is sacrificed in favor of stability and in most cases UCAVs do not fare well

when compared with classical transport airplanes.

Active Flow Control (AFC) may provide a solution to this conundrum by raising the trimmed lift coefficient

of an inherently unstable tailless aircraft configuration like the SACCON model [4]. SACCON is an acronym

describing the primary mission of the model: Stability And Control CONfiguration, because it was specifically

designed to raise the awareness of the aerodynamic research community to the stability and control issues of

a tailless Unmanned Combat Air Vehicle (UCAV). The aerodynamic characteristics of the SACCON model

and a description of the flow over it were documented by Vallespin et al. [5] while an assessment of the

predictive methods used is given by Cummings et al [6]. One could maintain the neutral point on the basic

model up to a lift coefficient of CL=0.25, hence by shifting its designed center of gravity it could presumably

be trimmed up to this value of CL.. However, the use of a small number of sweeping jet actuators enabled one

to trim the model up to CL=0.6, or more than double its potentially useful lift coefficient [4].

Experiments carried out on additional wind tunnel models shown in Figure 2 provided similar improvements

in longitudinal stability of these models [7]. Although the models differed from one another, they all resided

close to the neutral stability curve. The AFC actuators were usually located near the hinge lines of the control

surfaces and the jets that they emitted were either orthogonal to the hinge or to the TE. Actuation near the LE

was mostly unsuccessful, while actuation at the aft portion of the wing could outperform a large deflection of

a conventional control surface.

In order to avoid the guesswork pertaining to the optimal location of the actuators, their quantity, distribution

and orientation, an optimization procedure needs to be undertaken and the first step is to provide a reliable

database for which the hierarchy of the many variables is established. Thus, a simple wing based on the very

popular NACA0012 airfoil was selected for the initial set of tests. The wing, having an AR of 5.15 and sweep

of Λ= Λc/4=45°, is considered to be highly unstable (Figure 2) and provides a challenge for AFC. The test were

carried out at incompressible speeds at Mn≈0.1 and at Re=U∞(c/cosΛ)/ν ≈6*105.



Paper ID: 294 5

3 Experimental Apparatus

Tests were carried out at the University of Arizona in the closed-loop subsonic research wind tunnel (Figure

3) which has a test section that is 3ft high, 4ft wide and 12ft long. The maximum tunnel speed exceeds 80 m/s

(or 160knots) providing a Reynolds number of approximately 2*106 per foot at Mach Number, Mn=0.23. The

mean flow uniformity across the test section outside the wall boundary layers was better than ±0.5%, while

the turbulence intensity measured by a linearized hot wire anemometer that was filtered between 1Hz to 10kHz

was 0.05%. The test section temperature was maintained at 72°±1°F throughout the entire test campaign. An

artificial wall was inserted next to the wind tunnel’s side-wall in order to minimize the effects of the side wall

boundary layer. All forces and moments were measured by a 5-component external pedestal balance and the

swept back semi-span wing tested was attached to the balance by means of a special mount that also passed

the compressed air supply to the actuators.

Figure 3: The U of A research tunnel, the test article, and a schlieren picture of a sweeping jet in action [8].

A NACA 0012 airfoil section [9] that was swept back at Λ=45° and made out of extruded aluminum was used

in the present tests. The wing has a chord of 8” measured normal to the leading edge and is not tapered, twisted

or flapped. It has an AR of 5.15 with its wing-tip being perpendicular to the LE. It is thus highly unstable

according to the criteria of Figure 2. In order to avoid the formation of laminar bubbles and the associated

strong Reynolds number dependence, a “zig-zag” tape was used for tripping the boundary layer on both, the

top and bottom surfaces. Two small inserts at the LE of the wing divided its span into three equal segments

and were equipped with sweeping jet actuators blowing downward and forward, thus emulating a LE vortilon

(blue in Figure 3). The sweeping motion described a plane that was normal to the surface. Two additional

actuators located on the aft part of the upper wing surface emitted jets whose sweeping motion was parallel to

the surface. Different plugs containing the actuators were printed, each ejecting a jet that was inclined at a

different angle to the TE (Figure 3, orange: upstream actuator, green: downstream actuator). Each of the four

actuators could be activated independently or in conjunction with others since the purpose of the test was to

determine whether a single actuator could affect substantially the pitch break.

Sweeping jet actuators were used since they do not just add momentum to the boundary layer, they also cover

a large surface area in addition to creating streamwise vortices. Although the sweeping jet actuators do so

without any moving parts, they are more difficult to build than ordinary jet nozzles and the pressure drop across

them is larger. Consequently, they put a larger load on the air supply compressor whose size & power



Paper ID: 294 6

consumption are determined by the required mass flow and the pressure rise. Nevertheless, by sweeping from

side to side at an included angle of 100° (Figure 3, [8]) they are less sensitive to the precise knowledge of the

surface flow direction with which they are supposed to interact. The actuators are 3D printed and are an integral

part of an interchangeable plug allowing the axis of the jet that they emit to vary at angles β = -45° (inboard

direction); 0° (normal to the TE); 45° (outboard direction) & 90° (parallel to the TE in the outboard direction).

Oil flow and tuft visualization were carried out in addition to the measured integral forces and moments, while

the detailed pressure distribution on the wing at various incidence angles, actuation levels, location and

orientation of the actuators was obtained by using Pressure Sensitive Paint (PSP). The procedure used to

obtain dimensionless pressure coefficient contours is described in some detail in reference [4] and will not be

repeated presently.

4 Discussion of Results

Longitudinal, static instability is synonymous to pitch-up departure with increasing incidence angle, or with

increasing CL provided ∂CL/∂𝛼 is constant. The correlation provided in Figure 2 relates AR and the sweepback

of the quarter chord, Λc/4, to that instability. The Mean Rotation Line (MRL) about which CLM is measured

intercepts the longitudinal axis where the average chord (c=S/b) intercepts Λc/4. The origin of this choice is

attributed to the thin airfoil theory where c/4 is also the “aerodynamic center” or the “Neutral Point” (NP) on

finite wings. The quarter chord may not necessarily coincide with the NP on tailless aircraft where other design

parameters may be significant. Two potentially valid MRL locations may be used on the present test article

about which CLM is constant over a limited range of incidence angles. It should also be mentioned that ∂CL/∂𝛼

≠ constant over the entire range of incidence angles prior to stall. When the LE of the wing was clean and

actuator cavities were covered, trimmed conditions (CLM=0) existed for 1°<α<6° when MRL =1.315c

corresponding to CL<0.35 (Figure 4 center). On the other hand another neutral point exists (where ∂CLM/∂𝛼

≈∂CLM/∂CL=0) when MRL=1.278c for higher angles of incidence: 4°<α<9°, corresponding to 0.25<CL<0.51

(Figure 4 right). Thus, a shift in the MRL of 3.27% of the chord changed the entire stability character of the

wing. Moderate pitch up departure for the larger MRL=1.315c starts at α=6° that peaks three degrees later

resulting in a ΔCLM=0.0065. At this precise incidence (α=9°) the results corresponding to the smaller MRL

=1.278c initiate a nose down CLM with increasing α attaining a minimum ΔCLM departure from the respective

neutral point of ΔCLM=-0.0128 at α≈12°. Both non-linear pitch departures result also in the change of the lift-

curve slope. At α>12° an unstable (nose-up) pitch departure occurs that attains a CLM=0.13 or 0.09 depending

on the MRL selected. One may first ask whether deviations of ΔCLM ≈0.01 from a neutral point are significant?

The reply is affirmative because very similar numbers were obtained on the SACCON model and the

nonlinearities observed on this model raised many concerns [4] [5] [6].

Figure 4: Lift and pitching moments as functions of α for two different MRLs



Paper ID: 294 7

Flow visualization using tufts and fluorescent oil suggested that there was no leading edge separation or LE

vortex formation up and including α=9°, and consequently there was no pitch down departure corresponding

to MRL=1.278c which is usually associated with the creation of the LE vortex. On the other hand, the initial

pitch up departure for MRL=1.315c started at lower α and reached its maximum excursion at α=9°. The lift

slope curve, ∂CL/∂𝛼, was reduced between 5°<𝛼<9° suggesting that separation took place at this range of

incidence angles that was confirmed by the oil flow (Figure 5 top right).

Figure 5: Tuft & oil flow visualization on the wing

Separation near the TE wing-tip region results in high pressure that causes the CLM nose-up pitch departure for

both MRLs. The nose down trend started at 𝛼>9° where a LE vortex (bubble) was formed. At slightly higher

incidence a LE vortex was formed outboard of the selected MRLs, lowering the pressure behind them and

generating nose down pitch departure accompanied by an increase in ∂CL/∂𝛼. The origin of the LE vortex

propagates inboard with increasing 𝛼 and it intercepts the MRLs around 𝛼≈11°. Further increase in incidence

lowers the pressure inboard of the MRL starting the nose-up pitch departure. At 𝛼=11° the LE vortex spans

the wing diagonally originating at mid span at the LE and departing the wing at the TE near its very tip. It

therefore interacts with the separated area at the TE reducing its influence at the TE tip region (Figure 5 center

left). The origin of the LE vortex moves rapidly inboard with increasing 𝛼 and almost reaches the wing root at

𝛼=12° (Figure 5 center right). Since the flow visualization was carried out on a semi-span model, there is a

necklace vortex at the root of the wing with opposite sign of rotation that prevents the movement of the LE

vortex to the very root. In fact there is very little movement inboard of the origin of the LE vortex when the

incidence is increased beyond 𝛼>12° (Figure 5 bottom). The big difference is in the trajectory of the LE vortex



Paper ID: 294 8

whose foot print on the surface moves inboard with increasing 𝛼, so that at 𝛼=18° the inboard side of the vortex

intercepts the TE at mid span and it presumably lifts off from the surface around that location. The region of

reversed flow near the tip of the TE spreads out upstream and inboard at 𝛼>12°, reaching the LE trip strip at

the tip of the wing at 𝛼>18° (Figure 5 bottom right).

Early on, an attempt was made to emulate a LE fence (Figure 6 right: a photograph of the F-84F wing and a

sketch of the sweeping jet installation) by using sweeping jet actuators that were activated simultaneously from

top and bottom surfaces, as did the fence in the enclosed photograph. The results were deleterious, but they

exposed the sensitivity of the LE vortex to the tiny cavities located on both surfaces of the wing. Taping over

these cavities (which were only 0.025” wide) changed the pressure distribution over the upper surface of the

wing at incidence angles, 10°<α< 14°. With the slots being taped there was a single LE vortex that turned

gently downstream near the mid span of the wing (Figure 6 left). Removal of the tape in the absence of

actuation generated three weaker vortices that separated from the LE near each of the actuator slots and near

the wing tip (Figure 6 center). In order to simulate a vortilon the plugs containing the actuators have been

changed to ones containing only a single actuator (0.05” wide) blowing downward and somewhat against the

free stream, consequently there were no cavities on the upper surface of the wing. In order to assess the

sensitivity of the wing to the LE slot cavities, four balance baseline runs were carried out. Both cavities were

initially taped over, then the inboard and outboard cavities were sequentially uncovered and finally both

cavities were open. The effects of exposing the cavities on the lift and the drag were miniscule but the effects

on the pitching moment were substantially larger, particularly when the outboard or both cavities were

uncovered. In this case, the early non-linear excursions shown in Figure 4 were eliminated by the exposed

cavities, suggesting that the interaction between the LE vortex and the expanding TE tip separation was

reduced or eliminated

The open LE slot cavities restricted the motion and the strength of the LE vortex (Figure 6) by creating a new

vortex at each cavity, thus preventing the nose-down pitch departure that occurred in their absence around

α>9°. The pitch polar (i.e. the plot of CL vs CLM) suggests that locating the MRL at x/c=1.278 extends the range

of incidence angles at which the NP is approximately maintained to α≈12°. The absolute value of CLM at the

NP is larger when both cavities are exposed (it is more negative) rather than a single cavity. The nose up pitch

break occurs at lower angle of incidence when either the outer cavity or both cavities are exposed to the free

stream, due to the reduced interaction with the LE vortex. The choice of the second MRL located at x/c=1.315

is inappropriate as it does not enable a definition of a NP (Figure 7 bottom right) in the absence of actuation.

Figure 6: LE fence on the F-84F, a fluidic fence attempting to emulate it & the effect of slots on the LE vortex



Paper ID: 294 9

Figure 7: Pitching moment measured for the baseline wing with & without the exposed cavities

One might mention the sensitivity of CLM to excrescences located near the LE of the wing, because the LE

cavities could be taped over using two Kapton tapes of thickness .001” as shown in Figure 6 or by a continuous

single tape along the entire span thus avoiding the creation of backward and forward facing steps encountered

by the spanwise flow near the wing’s attachment line. The differences in CL are indistinguishable but not in

CLM (Figure 8).

Figure 8: The effect of tape discontinuity at the LE on CLM

When the LE is smooth the wing can be trimmed up to CL=0.28 or up to α=5° (Figure 9). Increasing α results

in nose up pitch break reaching a maximum CLM=+0.075 corresponding to α=9°. This pitch break can be

avoided by using a single actuator of the two shown in Figure 3. When the axis of the sweeping jet is normal

to the LE the aft actuator is more effective providing the ability to trim the wing up to CL=0.7 without exceeding

the input level of Cμ=0.26%. One may reach trimmed conditions up to CL=0.75 by increasing the available

input of momentum to Cμ=0.40%. Using both actuators at identical aggregate Cμ had only a deleterious effect

and doubling the Cμ did not increase the lift under trimmed conditions.



Paper ID: 294 10

Figure 9: The use of AFC emanating from the downstream or upstream actuator, or both at various input Cμ levels.

In order to assess the significance of the sweeping jet orientation, individual plugs containing actuators of

different inclination angles relative to the normal chord were made. This enabled AFC tests to be conducted at

45° intervals relative to the normal chord line (i.e. at β = 0°; ±45° & 90°). The results plotted in Figure 10

compare the efficacy of the actuation for the different actuation orientations relative to the normal chord when

only the downstream actuator was used. The momentum input was not changed throughout the experiment and

the data presented correspond to Cμ=0.26%. A ±45° degree deviation in either direction from the normal had

little effect on the CLM and on the NP that could now be maintained between 0.1<CL<0.7 instead of the short

range between 0.1<CL< 0.28 in absence of actuation. The most effective orientation over most of the incidence

angles considered is at +45° outboard. The 90° outboard orientation curve is almost parallel to the baseline

CLM results, except that it is shifted toward lower CLM values. It thus seems to affect the flow differently from

all other inclination angles. The 45° inboard orientation is the only one that had visibly lowered the lift at

α>11°.

Figure 10: The effect of single jet orientation on CL & CLM at a constant Cμ=0.26%.



Paper ID: 294 11

Since the aim of the current experiment is to assess the efficacy of AFC on the flow, the effects of the LE

excrescences (e.g. actuator cavities) have to be included. Exposing the two LE actuators but leaving them idle

brings up the question of the NP. Retaining MRL=1.278c defined a NP between 5° <α< 11° in the absence of

AFC. However, a single actuator that is inclined to the normal chord in the range of β = ±45° and uses the

lowest measurable Cμ input made the wing statically stable up to α=9° with an immediate pitch-up departure

at higher incidence angles (Figure 11 left). Therefore, if such actuation were to be used at angles of ±45° to

the normal chord in the presence of LE excrescences, MRL=1.315c should be selected for α>4° (Figure 11

right). The NP can be maintained in this case up to CL≈0.5 or a 9° incidence. This result would enable doubling

the usable lift coefficient of such a wing by using a single sweeping jet actuator at the lowest level of

momentum input.

There is however, a notable exception that would allow the introduction of AFC at 9° <α< 11° and still maintain

a NP up to CL≈0.75. It occurs when the sweeping jet axis is parallel to the LE blowing in the outboard direction

(see Figure 11 when β= 90°). AFC in this case should be used in conjunction of MRL =1.278c.

Figure 11: The effect of AFC on CLM with LE actuator cavities exposed to free stream (no actuation from LE)

The significance of the actuation location for the selected jet orientations and a prescribed Cμ=0.26% are

assessed by comparing the results shown already in Figure 11 and replotted presently onto Figure 12. The LE

cavities were taped-over in both cases, so the comparison involves the effect of location only. Actuation from

the downstream actuator at β= 45° generated CLM =-0.0069 in the range of 0.1<CL<0.4 while the same actuation

emanating from the upstream actuator generated CLM =-0.0053 and did so over a smaller range of incidence

angles. There is also a crossover between inboard blowing actuators and outboard blowing ones depending on

the range of α considered. The most obvious difference between the two figures corresponds to β= 90° that

generates a ΔCLM of 0.004 just due to the change in location. The downstream actuator is generally more

effective than the upstream one with the exception of β=90° at α>10°. The higher effectiveness of the upstream

actuator at β=90° is increased by having the cavities exposed to the free stream. This is attributed to the fact

that at β=90° the sweeping jet energizes the LE vortex when it is close to it and the LE excrescences keep the

LE vortices closer to the LE by preventing the LE vortex to become large and align itself more toward the free

stream in the outboard region of the wing. These differences are seen in the PSP photograph shown in Figure

6.



Paper ID: 294 12

Figure 12: The dependence of CLM on actuation location, orientation and input level

Visualization by tufts was also carried out on another wing based on the same extruded profile that had a

slightly larger AR of 5.68 and an array of sweeping jet actuators located at 75% of its chord. At α=9.3° there

was no indication of a LE vortex forming on this wing (Figure 13 top left) as there was no indication of it at

α=9° on the 5.15 AR wing (Figure 5). However, the last two rows of tufts indicated that the flow was parallel

to the TE over the outboard 2/3 of the span and on the verge of separation or slightly separated. This picture is

in complete agreement with the one shown in Figure 5. At α=9.4° (only a 0.1° higher than the previous picture

shown) a long LE vortex dominated the flow over the wing. Its origin was located between 35-40% of the span

(green tufts pointing outboard and upstream) and it covered a large triangular section extending to 50% of the

chord at approximately 70% of the span. Tufts located near the LE at the tip point in the direction of streaming

suggesting that the LE vortex merged with or became the tip vortex at mid chord near the tip. The sweeping

jets (axes normal to the LE) eliminated the LE vortex entirely at this incidence (magenta colored tufts) by

using a Cμ =0.5%. They must have altered the pressure gradient near the LE in order to achieve this large

effect. Similar observations were made at α=9.6° except that the origin of the LE vortex moved slightly

inboard. It is interesting to note that the outboard jets turned in the spanwise direction and entrain surface flow

toward them from the TE region (magenta tufts), thus accelerating the separation from this region. The curved

jet trajectory of the inboard actuation is clearly visible on this figure suggesting that the jet axis is

approximately aligned with the unperturbed surface flow and it seems to have no effect on the direction of the

flow upstream and inboard of its location (Figure 13). At α=10.7° the current level of the actuation no longer

prevented the generation of the LE vortex, but affected only the location of its origin as well as its inboard

extent. There is now clear upstream flow in the TE tip zone that is redirected downstream by the sweeping jets

(Figure 13 bottom right). This aspect of the baseline flow may be seen more clearly by the oil flow shown in

Figure 5.



Paper ID: 294 13

Figure 13: Superposition of tufts with (magenta) and without (green) AFC emanating from 2 pairs of actuators marked by magenta dots & the inboard location of the origin of the LE vortex as a function of α. (white tufts imply no effect of AFC or no AFC) [11]



Paper ID: 294 14

Although the main purpose of this investigation was to use AFC for control of pitch departure there are other

attributes of AFC that might be of interest. They will be discussed in detail for the open cavities case (i.e.

where the tiny cavities are considered as excrescences) while only the downstream actuator was used.

However, the effects of other aspects (such as a smooth LE or a change to an upstream actuator) is mentioned.

Potential control of yaw is of significance during cruise and loiter for military applications where the use of

regular control surfaces such as spoilers or split flaps is to be avoided [10] [7]. Since a semi-span model was

tested, the results measured in the absence of AFC (baseline) were subtracted from the actuated data assuming

that a complete wing is symmetrical about its longitudinal axis. Since incidence angles during cruise and loiter

are small the data provided will be limited to α<10° corresponding to CL≈0.5.

Figure 14: CLN; CLL; CLM as functions of CL (Top) when a single upstream actuator was active at Cμ=0.26%. CL /(CD + Cμ) vs. CL & CL vs. α (Center). The coupling between CLM & Δ CLN (Bottom)

The best yaw response for actuation emanating from the upstream sweeping jet at a Cμ =0.26% corresponds to

β=0° but even the downstream sweeping jet that is preferred for pitch control generates a yaw coefficient of

ΔCLN=-0.01 around CL≈0.5 (Figure 14 top left) and about half as much at incidence corresponding to CL=0.35.

The rolling moment, ΔCLL for the same vales of CL is somewhat smaller. By assuming that one may trim the

wing between 0.1<CL<0.5 using this actuation, one has to determine the coupling between CLN & CLM in order

not to lose pitch control while attempting to yaw in cruise or loiter. The cross plot of yaw vs. pitch coefficients

provides the needed information (Figure 14 bottom) suggesting that one may obtain the necessary yaw without

the need to adjust the pitch. These results compared very favorably to ΔCLN obtained on the SACCON or

MAGMA tailless aircraft models. These models required much larger momentum input coming out from

multiple actuators in order to obtain similar ΔCLN at comparable CL. Taping the LE and increasing the Cμ input

had a large effect on the yawing moment generated in this manner. The effectiveness of the current

configuration is attributed to the large AR of the model that provided a large moment arm although the current



Paper ID: 294 15

actuation position was not optimized for this purpose. A ΔCLN=-0.04 was easily attained when both actuators

were used at Cμ=0.8%. In the latter case the ratio of ΔCLN/ ΔCLL≈3 instead of the usual ratio of 1/3.

The CL/CD ratio of the basic wing is 12 in spite of the fact that its upper surface was rough because it was

painted with PSP (pressure sensitive paint) and relatively high trip strips both surfaces. Actuation increases

CL/CD but it often reduces CL/(CD+Cμ) and this case is an exception provided β=45° (Figure 14 center left).

However, the reduction in the maximum CL/(CD+Cμ)max is not large for all other cases considered and it depends

on the angle β and the location of the sweeping jet actuator. In fact when the upstream actuator was used at Cμ

=0.26%, the CL/(CD+Cμ) ratio was also somewhat improved for CL>0.5 suggesting that the entire momentum

input can be recovered as thrust.

4 Concluding Remarks

Tests were carried out on a 45° swept back simple wing of high aspect ratio that according to preliminary

design data sheets is highly unstable in pitch. The wing being based on a NACA0012 unswept airfoil had a

round leading edge. The airfoil is considered a leading edge staller, even though it is a border like case between

sharp leading edge airfoils and thick rounded ones. It has non-linear pitch characteristics that are either stable

or unstable depending on the chord location around which the moment is calculated. In short, the sweep back

of the ¼ chord whose selection is rooted in thin airfoil theory is no longer the determining factor in the

longitudinal stability of this wing. The nose up pitch departure starts at a low incidence angle of 5° prior to the

formation of a LE vortex and it is attributed to the reverse flow initiated near the tip of the wing at its TE. The

LE vortex, that is formed at α>9° close to the mid-span of the wing, aligns itself with the free stream direction

thus crossing the wing surface diagonally toward the tip of the wing. It interacts with the reverse flow observed

at lower incidence thus reducing its area of influence and stabilizing the wing in pitch. At higher α the LE

vortex, that turned into a wing tip vortex, lifts off from the surface enabling the resurgence of the reverse flow

and a secondary pitch-up instability that is no longer reversible prior to a complete stall of the flow. The

complex interaction between the LE vortex and the TE separated flow can be easily affected by AFC. A single

jet can change the orientation of the LE vortex or change its inner strength (circulation) depending on the

injected momentum, the location of its origin and its direction. Some of these effects were tested presently.

Placement of an excrescence or any discontinuity (a LE snag, a vortilon, a small fence or a cavity) at the LE

of a swept back wing has a major effect on the LE vortex and it is widely used to prevent wing-drop. It may

also help AFC by redirecting the LE vortex to a convenient location where an actuator is present. This happened

on the SACCON model where a sudden change in the LE radius of the wing caused the LE vortex to depart

the LE and turn in the spanwise direction in the aft part of the wing just upstream of the TE [7]. It provides a

good flow-physics-based example for the need to integrate AFC into a wing design procedure. The present

experiment suggests that it is possible to keep the wing trimmed by an additional 5° to 7° of incidence thus

enable it to double its usable lift coefficient. This was achieved using a single actuator at very modest input of

momentum.



Paper ID: 294 16

References

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wings," IHS ESDU, 2001.

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and J. Ladd, "On The Use of Active Flow Control (AFC) on Tailless Aircraft Models to

Affect their Trim and Control," in AIAA Scitech 2019, San Diego, California, 2019-0045.

[8] D. Hirsch, "An Experimental and Theoretical Study of Active Flow Control," Ph.D

Thesis, California Institute of Technology, 2017.

[9] I. Abott, A. von Doenhoff and L. Stivers, "Summary of Airfoil Data," NACA TR 824,

1945.

[10] S. Hitzel, "Perform & Survive, Evolution of Some U(M)CAV Platform Requirements,"

STO-MP-AVT-2015, 2015.

[11] M. Menge, Investigation of a Swept-Back Wing with Variation of Aspect Ratio using

Surface Flow Visializiation and Pressure Sensitive Paint, B.Sc. Thesis: Universitaet der

Bundeswehr Muenchen, 2017.

using sweeping jets on swept wings - unina.it · tests were carried out at the university of...

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