A vortex attenuation device is presented which dissipates a lift-induced vortex generated by a lifting aircraft wing. The device consists of a positive pressure gradient producing means in the form of a compression panel attached to the lower surface of the wing and facing perpendicular to the airflow across the wing. The panel is located between the midpoint of the local wing cord and the trailing edge in the chord-wise direction and at a point which is approximately 55 percent of the wing span as measured from the fuselage center line in the spanwise direction. When deployed in flight, this panel produces a positive pressure gradient aligned with the final roll-up of the total vortex system which interrupts the axial flow in the vortex core and causes the vortex to collapse

Patterson, James C., Jr.

English

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US005230486A 
United States Patent 1193 1111 Patent Number: 5,230,486 
Patterson, Jr. [45] Date of Patent: Jul. 27, 1993 
UNDERWING COMPRESSION VORTEX 
ATTENUATION DEVICE 
Inventor: 
Assignee: 
James C. Patterson, Jr., Newport 
News, Va. 
The United States of America as 
represented by the Administrator of 
the National Aeronautics and Space 
Administration, Washington, D.C. 
Appl. No.: 886,998 
Filed: May 22, 1992 
Int. C l . 5  .............................................. B64C 23/06 
U.S. C1. ..................................... 244/199; 244/198 
Field of Search ................ 244/198, 197, 200, 130 
References Cited 
U.S. PATENT DOCUMENTS 
OTHER PUBLICATIONS 
James Scheiman et al. “Exploratory Investigation of 
Factors Affecting the Wing Tip Vortex” Apr. 1972. 
Primary Examiner-David M. Mitchell 
Assistant Examiner-Christopher P. Ellis 
Attorney, Agent, or Firm-Kimberly A. Chasteen 
ABSTRACT 
A vortex attenuation device is presented which dissi- 
pates a lift-induced vortex generated by a lifting aircraft 
wing. The device consists of a compression panel at- 
tached to the lower surface of the wing and facing 
perpendicular to the airflow across the wing. The panel 
is located between the midpoint of the local wing cord 
and the trailing edge in the chord-wise direction and at 
a point which is approximately 55% of the wing span as 
measured form the fuselage center line in the manwise 
P71 
L 
3,984,070 10/1976 Patterson, Jr. . 
3,991,132 12/1976 Erwin . 
4,371,267 3/1983 Haworth ............................. 244/199 
4,955,958 9/1990 Dellinger et al. . 
5,037,044 8/1991 Seyfang ............................... 244/199 
direction. When deployed in flight, this panel produces 
a positive pressure gradient aligned with the final roll- 
up of the total vortex system which interrupts the axial 
flow in the vortex core and causes the vortex to col- 
lapse. 
FOREIGN PATENT DOCUMENTS 
0470287 5/1969 Switzerland ........................ 244/199 4 Claims, 1 Drawing Sheet 
14 
https://ntrs.nasa.gov/search.jsp?R=19940006218 2020-06-17T23:34:59+00:00Z
U.S. Patent July 27, 1993 5,230,486 
14 
FIG. 1 
50 % 
FIG. 2 
5,230,486 
1 2 
by a lifting aircraft wing having a leading edge, a trail- 
ing edge, an upper surface and a lower surface. The 
wing includes a positive pressure gradient producing 
means in the form of a compression panel attached to 
5 the lower surface of the wing and facing perpendicular 
UNDERWING COMPRESSION VORTEX 
ATTENUATION DEVICE 
ORIGIN OF THE INVENTION 
~. 
The invention described herein was made by employ- 
ees of the United States Government and may be used 
by and for the Government for governmental purposes 
without the payment of any royalties thereon or there- 
for. 
BACKGROUND OF THE INVENTION 
to the airflow across the wing. The panel -is located 
between the midpoint of the local wing cord and the 
trailing edge in the chord-wise direction and at a point 
which is approximately 55% of the wing span as mea- 
10 sured from the fuselage center line in the spanwise di- 
rection. The panel is hinged so that is may be retracted 
when its use is unnecessary. This apparatus does not 
The present invention relates generally to aircraft 
aerodynamics and specifically to a method and appara- 
tus for attenuating the lift-induced vortex generated by l5 
the wings of an aircraft. 
provide elevational control; its purpose is to dissipate 
the total vortex system of an aircraft. 
BRIEF DESCRIPTION OF THE DRAWINGS 
The lift-induced vortex system associated with large 
jet aircraft has become a major problem to the air traffic 
controller as an unseen hazard to smaller following 
aircraft. The problem arises from the fact that the vor- 20 
tex system created by large aircraft is very strong and of DETAILED DESCRIPTION OF THE 
such long persistence and that it intensifies as the weight INVENTION 
of the aircraft increases. The rotational airflow or vor- FIGS. 1 and 2 show an aircraft wing 12 extending 
tex developed at or near the wingtip of an aircraft re- from a fuselage 14. A compression panel 16 extends 
mains essentially stationary in space and induces a roll- 25 from the lower surface of the wing 12. The preferred 
ing moment on smaller following aircraft is encoun- location of the panel 16 is between the midpoint and 
tered. This rolling moment can be so strong that it sur- trailing edge ofthe wing 12 in the wing chord direction 
passes the limits of control of the smaller aircraft and and at a point which is appro~mately 55% of the wing- 
such loss Of has been proven to Occur as far as span as measured for the fuselage center line. The sur- 
4 miles behind a large aircraft. The vortex problem is 30 face area of panel 16 is approximately one half of one 
where smaller aircraft are most likely to follow large for the panel 16 is rectangular with the width being two aircraft and during landing where the small aircraft has to three times the height. The size, shape and position of little altitude in which to regain control. 
The vortex system also causes the induced drag ofthe 35 the pane' l6 may vary depending On the aircraft and the 
FIG. 1 is a front view of an aircraft wing; and 
FIG. 2 is a cross sectional view taken along line 
11-11 FIG. 1. 
particularly in flight paths around major percent of the total wing surface area. The ideal shape 
aircraft which, for transport-type aircraft, is 35 to 40% 
signed to shed the vortex system along the span of the 
wing to reduce the concentrated vortex centered just 
flap system 
pressure gradient 
On that aircraft. The panel l6 be 
of the total drag of the aircraft. Aircraft wings are de- 
behind each wingtip. The final roll-up of the vortex 40 the vortex core and causes the vortex to collapse. The 
system is further behind the wing thereby reducing, but Panel 16 is hinged SO that it may be removed f r m  the 
not eliminating, its ill effect on the drag of the aircraft. airstream when not in use. The Panel 16 Will be bwered 
Previous attempts to limit the vortex effect have into a position (see arrow 18) which is perpendicular to 
generally been unsuccessful either because, as with the airflow when the aircraft is likely to encounter other 
wing end plates or wingtip extensions, the devices did 45 aircraft which would be disturbed by its vortices, such 
not sufficiently modify the far downstream effects of as in takeoff and landing. 
the vortex or, as with landing flap variations and wing The placement of the panel 16 is critical to ensure that 
spoilers, the vortex attenuation caused too great a lift the entire vortex System is dissipated. At the position 
loss for the aircraft. These devices are only concerned indkated above, the Panel 16 Will attenuate the wingtip 
with attenuating wingtip vortices, not with attenuating 50 Vortex and the Smaller Vortices created along the entire 
the entire vortex system. wingspan. 
It is an object of the present invention to attenuate the What is claimed is: 
total vortex system produced by an aircraft. 1. An apparatus for dissipating a lift-induced vortex 
It is another object of this invention to provide an System generated by an aircraft Wing in an airflow hav- 
apparatus which is effective on all types of aircraft. 55 ing a leading edge, a trailing edge, an upper surface and 
It is a further object of this invention to provide an a bwer surface comprising: a positive pressure gradient 
apparatus which causes an increase in wing lift rather producing means; said gradient producing means being 
than a reduction when deployed. a panel which is attached to the lower surface of the 
It is yet another object of the present invention to wing and which faces in a direction perpendicular to 
accomplish the foregoing objects in a simple manner. 60 the airflow, wherein said panel is located such that it 
Additional objects and advantages of the present produces a positive pressure gradient aligned with the 
invention are apparent from the drawings and specifica- final roll-up of the total vortex system which interrupts 
tion which follow. the axial flow in the vortex core and causes the vortex 
consiructed and located such that is produces a positive 
with the final roll-up of the. 
vortex system which interrupts the flow in 
to collapse, said location being approximately 55 per- 
65 cent of the wing span as measured from the fuselage 
center line and between the midpoint of the local wing 
chord and the trailing edge, whereby the vortex system 
is eliminated. 
SUMMARY OF THE INVENTION 
According to the present invention, the foregoing 
and additional objects are obtained by providing an 
apparatus for dissipating a lift-induced vortex generated 
5,230,486 
3 4 
2. The apparatus of claim 1 in which the panel is produces a positive pressure gradient aligned with the 
hingedly attached to the lower surface of the wing. final roll-up of the total vortex system which interrupts 
3. An apparatus for dissipating a lift-induced vortex the axial flow in the vortex core and causes the vortex 
generated by an aircraft wing in an airflow comprising to collapse, said location being approximately 55 per- 
an aircraft having an elevator control device, a fuselage 5 cent of the wing span as measured from the fuselage 
wing having a leading edge, a trailing edge, an upper center line and between the midpoint of the local wing 
surface, a lower surface and a positive pressure gradient chord and the trailing edge, whereby the vortex system 
producing means; said gradient producing means being is eliminated. 
a panel which is attached to the lower surface of the 4..The apparatus’of claim 3 in which the panel is 
wing and which faces in a direction perpendicular to 10 hingedly attached to the lower surface of the wing. 
the airflow, wherein said panel is located such that it * * * * *  
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Underwing compression vortex attenuation device

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