Some airfoil designs have been shown by theory and small-scale tests to be capable of passively maintaining laminar flow at super -sonic speeds. More recently, flight tests have proven that these designs can maintain large runs of laminar flow at higher Reynolds numbers in harsh flight environments. The flight tests were conducted for the purposes of observing laminar flow at speeds up to mach 2.0 and determining the conditions under which laminar flow breaks down.

The Boundary of the Transition from laminar to turbulent flow is visible in an infrared image of the test article taken during supersonic flight.
Flight tests were performed on an F- 15B airplane equipped with a laminar- flow test article — essentially, a small half wing mounted vertically (see figure). The airplane was also equipped with an infrared camera for ob -serving the test article flow. In effect, the camera measures the local surface temperature of the test article, which temperature varies with the state of the boundary-layer flow: The portion of the surface covered by a turbulent boundary layer is warmer than that covered by a laminar boundary layer — a consequence of its higher wall recovery temperature as well as greater con -vection of the turbulent layer with the free-stream flow. Hence, in the infrared image, the surface in the turbulent-flow region appears brighter than that in the laminar-flow region.

The test article was fabricated from aluminum with an insulating layer covering all but the first 3 to 4 in. (about 7.5 to 10 cm) in from the leading and trailing edges. A splitter plate was installed at the root of the test article to minimize the effect of disturbances from the bottom of the airplane on the flow about the test article and to better simulate a full-span wing.

Laminar flow was observed up to full chord on the outer third of the span of the test article and up to approximately 80 percent of chord over the inner two-thirds of the span. The laminar flow was found to be capable of penetrating weak shock waves, but was typically terminated by strong shock waves. The strongest shock wave incident on the test surface appeared to emanate from the camera pod, which was located on an armament rail of the airplane.

For future flights, the test article will be instrumented with surface-pressure gauges and thermocouples to obtain more detailed data. Also, the video data-recording system will be updated to enable it to record the full 12-bit digital images from the camera. A contemplated subsequent program would use a larger test article that would permit assessment of effects of greater Reynolds numbers.

This work was done by Daniel W. Banks of Dryden Flight Research Center and Richard R. Tracy and James D. Chase of Reno Aeronautical Corporation, a subsidiary of Directed Technologies, Inc. Further information can be obtained from Daniel Banks [telephone no. (661) 276-2921 or e-mail This email address is being protected from spambots. You need JavaScript enabled to view it.].

In accordance with Public Law 96-517, the contractor has elected to retain title to this invention. Inquiries concerning rights for its commercial use should be addressed to Richard R. Tracy, Ph.D., President Reno Aeronautical Corporation 3000 Old Ranch Rd. Carson City, NV 89704-9542 Tel. No.: 775-883-8410 E-mail: This email address is being protected from spambots. You need JavaScript enabled to view it. Refer to DRC-00-22, volume and number of this NASA Tech

NASA Tech Briefs Magazine

This article first appeared in the June, 2002 issue of NASA Tech Briefs Magazine.

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