Today's modem fighter design, with multifimction nozzles, places an increased emphasis on nozzle/airframe integration. The tools currently available to the aircraft designer for aft-end design and evaluation are model test reports, being disseminated mainly by government laboratories, and three-dimensional numerical computation codes. Test data utilization usually is limited by the suitability of the area that has been tested. The second approach, analysis, usually requires timeconsuming three-dimensional configuration data input. Recognizing the need for a quicker means of solution, useful in a preliminary design environment, a semi-empirical computer methodology has been developed for determining threedimensional aircraft afterbody performance. The essence of the approach is to construct equivalent bodies of revolution of three-dimensional bodies and then to utilize a straight or hybrid axisymmetric analysis. This approach was developed for single-and twin-engine axi symmetric and two-dimensional afterbodies. The computer code covers the Mach number range of 0 to 3.5, and includes boundary layer flow and plume entrainment calculations. The methodology was verified by comparing afterbody drag and axial and longitudinal pressure distributions. The isolated drag coefficients are then modified to account for three-dimensional afterbody flow field effects and also for aircraft component effects such as empennage, booms, interfairings, base areas, and spacing. NomenclatureS/D spacing ratio (spacing between engine/nozzle dia. at customer AR aspect ratio, area ratio connect) A/B afterburner Τ temperature ATS air-to-surface T/C thickness to chord ratio CD drag coefficient V