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NOTICE

NASA CASE NO. LAR 14470-1

PRINT FIG. 1

The invention disclosed in this document resulted from research in aeronautical and

space activities performed under programs of the National Aeronautics and Space

Administration. The invention is owned by NASA and is, therefore, available for licensing

in accordance with the NASA Patent Licensing Regulation (14 Code of Federal

Regulations 1245.2).

To encourage commercial utilization of NASA-owned inventions, it is NASA policy to grant

licenses to commercial concerns. Although NASA encourages nonexclusive licensing to

promote competition and achieve the widest possible utilization, NASA will provide the

necessary incentive to the licensee to achieve early practical application of the invention.

Address inquiries and all applications for license for this invention to NASA Patent

Counsel, Langley Research Center, Code 143, Hampton, Virginia 23665-5225.

Serial No.: 07/823,809

01-07-92 LaRC

(NASA-Case-LAR-14470-t) BOUNDARY

LAYFR RELAMI_A_[ZAT[0N DEVICE

Patent Application (NASA) 13 p

N93-I1876

Unclas

G3/O2 0123664
BOUNDARY LAYER RELAMINARIZATION DEVICE 

AWARDS ABSTRACT 

NASA Case No. LAR 14470-1 

A need has existed for a device to manipulate flow on the leading edge of a swept wing 

by artificial means at supersonic speeds, in order to present a low drag configuration. It 

has been known that by using suction, the flow over a wing can be relaminarized and 

then maintained laminar over 75 percent of chord length. However, suction is an active 

device and vacuum pumps, tubing and necessary switches and valves are required to 

make such a system work. Such pumps, tubes, switches, and valves add weight to the 

aircraft, and require power that a supersonic aircraft could better use somewhere else. 

Moreover, suction devices are hard to fit in the small leading edge of most present day 

supersonic aircraft. Another negative factor is the high manufacturing cost of the suction 

system. 

According to the present invention, relaminarization of a boundary layer formed in 

supersonic flow over the leading edge of a swept airfoil is accomplished by means of at 

least one band, especially a quadrangular band, and most preferably a square band. 

Each band conforms to the leading edge and the upper and lower surfaces of the swept 

airfoil as an integral part thereof and extends perpendicularly from the leading edge. 

Each band has a height of about two times the thickness of the maximum expected 

boundary layer. 

Special advantages are that the passive devices according to the present invention are 

highly efficacious, at the same time very simple and quite easy to use. They may be 

fabricated as ribs in the surface of the airfoil itself, or they may be machined separately 

and retrofitted to any existing supersonic aircraft. 

Inventor: Theodore R. Creel 

Address: 206 York Warwick Drive 

Employer: NASA LaRC 

Initial Evaluator: Dennis M. Bushnell 

Serial No.: 071823,809 

Filed: January 7, 1991 
I

Serial No. 07/823,809

Filed: 01-07-92

LAR 14470-1 -1- PATENT APPLICATION

BOUNDARY LAYER RELAMINARIZATION DEVICE

Ori.(:lin of the Invention

5 The invention described herein was made by an employee of the United

States Government and may be manufactured and used by or for the

Government for governmental purposes without the payment of any royalties

thereon or therefor.

10 Backqround of the Invention

o

m Field of the Invention

15

This invention relates generally to the reduction of total airplane drag.

In particular, it relates to reducing total airplane drag by changing high drag

turbulent wing leading edge flow to low drag laminar wing leading edge flow in

supersonic flight regimes.

20

25

30

.

u Description of Related Art

The destabilizing effects of leading-edge sweep on the laminar wing

boundary layer at subsonic speeds have been known since the early 1950s.

The dominant transition mechanisms are "attachment-line contaminations" and

"crossflow instability." The problem of turbulent contamination along the

attachment line of swept wings had to be addressed in the 1960s before

laminar flow could be achieved at high Reynolds numbers on swept wings.

(See, Pfenninger, W.: Flow Phenomena at the Leading Edge of Swept Wings.

Recent Developments in Boundary Layer Research - Part IV, AGARDograph

97, May 1965; Pfenninger, W.: Laminar Flow ControI-Laminarization,

AGARD-R-654, March 1977, pp. 3-1 - 3-75; and Gaster, M.: On the Flow
LAR 14470-1 -2- PATENT APPLICATION

Along Swept Leading Edges. The Aeronautical Quarterly, Vol. XVIII, Part 2,

May 1967, pp. 165-184.) In some of these studies, leading edge contamination

was minimized by strong local suction and leading-edge fences with suction.

In others, a faired bump protruding from the leading edge of a wind-tunnel

5 model wing was used to prevent the spanwise propagation of turbulence along

the attachment line. Wind tunnel tests of this bump attached to a cylindrical

model indicated that for subsonic flight, laminar flow could be maintained at

sweep angles of 60°.

Evidence of crossflow instability was first observed in flight experiments

10 on the leading-edge portions of swept wings as regularly spaced streaks

caused by variations in mass transfer from surface coatings. Gray, W.E.: The

Nature of the Boundary-Layer Flow at the Nose of a Swept Wing. RAE TM

Aero 256, 1952. The streaks were aligned in the local streamwise direction and

were present over the upstream region preceding transition on swept-back

15 wings. These streaks were caused by co-rotating vortices arising from the

inflectional instability of the crossflow boundary-layer profiles in the upstream

region of the swept wings. Gregory, N.; Stuart, J.T.; and Walker, W.S.: On

the Stability of Three-Dimensional Boundary Layers with Application to the Flow

Due to a Rotating Disk. Philos. Trans. Roy. Soc., London, Ser. A, Vol. 248,No.

20 943, 1955, pp. 155-199.

In flight, spanwise contamination may be minimized by either active or

passive means. The cheapest method for avoiding turbulent attachment-line

flow would be a passive device to either prevent turbulence from spreading

down the leading edge, or to relaminarize an already turbulent boundary layer.

25 Recent subsonic experimental evidence suggests several passive devices to

relaminarize leading edge-flow in the subsonic region. Seyfang, G.R.:

Turbulent Drag Reduction by Passive Means. Proceedings of the International

Conference, London, England, Sept. 15-17, 1987, Vol. 2, London, Roy. Aero.

Soc., 1987, pp. 568-601. Seyfang's wind tunnel results for a swept-wing model

30 covered sweep angles of 25° to 75° at M, = 0.25.
5

LAR 14470-1 -3- PATENT APPLICATION

That such devices should have utility in the supersonic region is not to

be expected by those of skill in the art, as it is well known that a device which

is useful in the subsonic region is often not useful in the supersonic region.

Moreover, it would appear that when operating in the supersonic region, a

boundary layer should become more turbulent as a result of employing such

devices.

Summary of the Invention

10

15

20

25

It is the primary object of the present invention to provide what has been

heretofore unobtainable in the prior art, viz., a device for manipulation of the

flow on the leading edge of a swept airfoil at supersonic speeds in order to

obtain a low drag configuration.

This object and its attending benefits are achieved by providing a band

which conforms to the leading edge and the upper and lower surfaces of the

swept airfoil as an integral part thereof, the band extending perpendicularly from

the leading edge of the swept airfoil and having a height of about two times the

thickness of the maximum expected boundary layer. Especially good results

are obtained if the band is quadrangular (especially a square band). The band

may be affixed to the swept airfoil by standard means (e.g., mechanical means

or cementing means), or the band may be formed as a rib in the surface of the

airfoil itself. Of course, a plurality of bands spaced from each other along the

spanwise length of the airfoil may be employed if desired. When the

supersonic flow is Mach 3.5 at a sweep angle of the swept airfoil of 76 o over

a freestream Reynolds number of 1.5 x 10s to 1.5 x 106, it is especially

advantageous if the height of the band (preferably square) is at least 0.01

inches, and preferably 0.02 inches.
LAR 14470-1

Brief Description of the Drawinqs

-4- PATENT APPLICATION

5

10

15

For a more complete understanding of the present invention, including

its object and attending benefits, reference should be made to the Description

of the Preferred Embodiments, which is set forth below. This description should

be read together with the accompanying drawings, wherein:

FIG. 1 is a schematic showing a preferred embodiment of a boundary

layer relaminarization device according to the present invention;

FIGS. 2 and 3 are side and top view representations of the boundary

layer surface flow effects produced by the boundary layer relamination device

of FIG. 1 acting on a swept cylinder model, as visualized using oil-flow

techniques; and

FIG. 4 is a schematic setting forth a hypothesis for the mechanism of

operation of boundary layer relaminarization devices according to the present

invention.

Description of the Preferred Embodiments

Referring now to the drawing, FIG. 1 shows a preferred embodiment of

20 a boundary layer relaminarization device according to the present invention,

which is a band 11 that conforms to the leading edge 12, as well as upper

surface 13 and the lower surface (not shown) of swept airfoil 14. Two such

bands 11 are shown in FIG. 1, wherein the direction of supersonic airflow is

shown by the arrow labeled 10. The boundary layer relaminarization device is

25 shown as a quadrangular band 11, especially a square band, wherein height

15 equals width 16. This embodiment achieved the most beneficial results.

The length 17 of band 11 may vary over a wide range without any beneficial

or deleterious effect upon the results. Of course, other configurations of band

11 may be employed with some success. However, the quadrangular,

30 especially square configuration is especially preferred. As shown in FIG. 1,
5

10

15

20

25

30

LAR 14470-1 -5- PATENT APPLICATION

band 11 extends perpendicularly from leading edge 12 of airfoil 14. As

determined empirically in supersonic flow, height 15 of band 11 must be about

two times the thickness of the maximum expected boundary layer to be

relaminarized, in order that effective relaminarization thereof can take place.

Devices which do not meet this specification are not efficacious.

Band 11 may be affixed to the leading edge 12 and upper 13 and lower

(not shown) surfaces of swept airfoil 14 by standard means, such as

mechanical means (riveting, bolting), or cementing means, or welding means,

so that band 11 becomes an integral part of swept airfoil 14. Alternatively,

band 11 may be formed as a rib in the surface of airfoil 14 in the original

fabrication thereof. It is often advantageous to employ a plurality of bands 11,

as shown in FIG. 1, which are spaced from each other along the spanwise

length of the airfoil. See also FIG. 4 and its accompanying discussion infra.

Band 11 may be fabricated from any material which can withstand the heat

generated by the supersonic flow, such as steel, other metal alloys known to

those of skill in the art, and high performance polymeric composites.

Employed initially was a swept cylinder model 18 (see FIGS. 2-4), which

consisted of an 0.030-inch stainless steel cylindrical shell of one inch outside

diameter with the upstream end sealed and cut off parallel to the freestream

flow at a sweep angle of 60°. In order to withstand the aerodynamic loads,

swept cylinder 18 was supported by a strong-back arrangement. Two rows of

0.01-inch diameter chromel-alumel thermocouple wires were spot-welded to the

inside surface of the cylinder shell, 80° apart, and at 0.25-inch intervals along

an 8.75 inch length of the shell, starting at 3.0 inches from the upstream tip of

the model. The 12-inch long swept cylinder was constructed so that the shell

could be rotated in 10° increments through an arc of e!-80°.

The surface of swept cylinder 18 was maintained clean and polished to

a finish of less than 10 rms micro-inches. Devices 11 were tested at 60° and

76° sweep over a freestream Reynolds number range of 1.0 x 105/inch to

1.88 x 108/inchin the NASA Langley Research Center Mach 3.5 Supersonic
LAR 14470-1 -6- PATENT APPLICATION

Low-Disturbance Pilot Tunnel.

To determine whether a swept cylinder boundary layer was laminar or

turbulent, computed recovery factors, as understood by those of skill in the art,

were compared to the theoretical laminar value of 0.85 and to the flat-plate

5 turbulent boundary layer value (r=0.89), as known to those of skill in the art.

Referring now to FIGS. 2 and 3, there are shown in side and top view

representations respectively the boundary layer surface flow effects produced

by a boundary layer relaminarization device 11 of FIG. 1, according to the

present invention, acting on swept cylinder model 18, as visualized using oil

10 flow techniques, as described hereinafter.

A mixture for oil-flow studies was prepared by mixing two parts by

volume of 350 centistokes silicon oil to one part titanium powder. A 1/4-inch

wide strip of the mixture was applied to the model 18 along the attachment line.

The model was then mounted in the tunnel and the desired stagnation pressure

15 was established as quickly as possible (two to three minutes) and held constant

for five to ten minutes, which was sufficient to stabilize the pattern over the

forward part of the model.

Representations of the typical oil-flow patterns on the cylinder 18 are

shown in FIGS. 2 and 3, where evidence of crossflow vortices 19 can be seen

20 downstream of square boundary-layer relaminarization device 11 having a

height of 0.5 inches. Upstream of device 11 the crossflow vortices have been

obliterated by the turbulent, boundary layer 20 at the test Reynolds number of

R_, g = 9 x 105.

Also shown in these figures are the boundary layer separation 21 and

25 reattachment 22 lines. At this Reynolds number, separation has occurred

approximately 0.28 inches upstream of the square device 11. After the

boundary layer passes over the square device 11, it reattaches approximately

0.09 inches downstream of the square device. Video recordings of the oil-flow

runs show that oil is flowing radially outwardly on both sides of device 11,

30 suggesting the formation of streamwise vortices at the square device 11.
5

10

15

20

LAR 14470-1 -7- PATENT APPLICATION

Laminar flow is seen at 23.

A qualitative hypothesis for the mechanism responsible for the success

of relaminarization devices according to the present invention can be

constructed from photographs and videotapes of the oil flow surface patterns

discussed above. A graphical representation of the hypothesized flow-field in

the vicinity of device 11 is shown in FIG. 4.

As turbulent attachment-line flow 24 on the leading edge approaches

device 11, the obstruction presented by device 11 forces the attachment-line

boundary layer to form trailing vortices 25 on either side of device 11. These

vortices 25 pumpthe turbulent attachment-line fluid "overboard," allowing a new

uncontaminated laminarattachment-line boundary layer 26 to form downstream

of device 11. This process may be repeated along the leading edge with

multiple devices 11, as depicted in FIG. 4.

As a result of actual testing, it has been shown that when the supersonic

flow is Mach 3.5 at a sweep angle of the swept airfoil model of 76° over a

freestream Reynolds number of 1.5 x 10sto 1.5 x 10_,the height 15of band 11

(especially square band 11) must be at least about 0.01 inches for

relaminarization to occur. Moreover, the very best results are achieved when

the height 15 of band 11 (especially square band 11) is at least about 0.02

inches, with values as high as 0.05 inches showing excellent results under

these conditions.
LAR 14470-1 PATENT APPLICATION

BOUNDARY LAYER RELAMINARIZATION DEVICE

Abstract of the Disclosure

5

10

Relamination of a boundary layer formed in supersonic flow over the

leading edge of a swept airfoil is accomplished by means of at least one band,

especially a quadrangular band, and most preferably a square band. Each

band conforms to the leading edge and the upper and lower surfaces of the

airfoil as an integral part thereof and extends perpendicularly from the leading

edge. Each band has a height of about two times the thickness of the

maximum expected boundary layer.
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