This paper describes a program of wind tunnel tests of the hybrid laminar flow control (HLFC) concept at near full scale Reynolds number. The tests were performed in the low speed 5′ X 8′ Boeing Research Wind Tunnel on a large (20 Foot Chord) section of an infinite swept wing having a sweep angle of 30°degrees. Boundary layer suction was provided over the first 20 percent chord on the wing upper surface through an electron beam perforated titanium suction surface. The extent of laminar run beyond the suction region on the wing surface was measured for various external pressure distributions and suction distributions. Two transition detection techniques were employed: an off-the-surface-pitot and a relocatable hot film. Depending upon the external pressure distribution the laminar run extended as far back as 45 percent chord. This corresponds to a transition Reynolds number of approximately 11 x 106. Significant spanwise non-uniformity of the transition location was encountered at “off-design” model incidences. This was caused by spanwise non-uniformity of suction induced by spanwise external pressure gradients. For given chordwise pressure distribution and Reynolds number conditions, the maximum chordwise extent of laminar run was found to be insensitive to the suction level over a wide range.Evidence of “aerodynamic roughness” induced transition was found under strong suction applied through multiple rows of discrete holes.