As part of a cooperative research program between NASA, McDonnell Douglas Corporation, and Wright Research and Development Center, a flow field investigation was conducted on a 7.52 percent scale windtunnel model of an advanced fighter aircraft design. The investigation was conducted in the Langley 16 ft Transonic Tunnel at Mach numbers of 0.6, 0.9, and 1.2. Angle of attack was varied from -4 degrees to 30 degrees and the model was tested at angles of sideslip of 0, 5, and -5 degrees. Data for the over the wing flow field were obtained at four axial survey stations by the use of six 5 hole conical probes mounted on a survey mechanism. The wing leading edge primary vortex exerted the greatest influence in terms of total pressure loss on the over the wing flow field in the area surveyed. A number of vortex control devices were also investigated. They included two different apex flaps, wing leading edge vortex flaps, and small large wing fences. The vortex flap and both apex flaps were beneficial in controlling the wing leading edge primary vortex. Bare, E. Ann and Reubush, David E. and Haddad, Raymond C. Langley Research Center DELTA WINGS; FIGHTER AIRCRAFT; FLOW DISTRIBUTION; SUBSONIC SPEED; TRANSONIC SPEED; VORTEX FLAPS; WIND TUNNEL MODELS; WIND TUNNEL TESTS; ANGLE OF ATTACK; FLOW VISUALIZATION; MACH NUMBER; PRESSURE DISTRIBUTION; SIDESLIP; TRANSONIC WIND TUNNELS...