Most fighters use the convergent-divergent nozzle configuration to accelerate into the supersonic realm. This nozzle configuration greatly increases the thrust potential of the aircraft compared to the simpler convergent nozzle. The nozzle design is not only crucial for thrust, but also for the drag since the afterbody drag can be as high as 15% of the total. Engine manufacturers optimize the engine and the nozzle configurations for the uninstalled conditions, but these may not be optimal when the engine is installed in the aircraft. The purpose of this study is to develop a methodology to optimize axisymmetric nozzle settings in order to maximize the net thrust. This was accomplished by combining both simulations of thrust and drag. The thrust model was created in an engine performance tool, called EVA, with the installed engine performance of a low bypass turbofan jet engine at maximum afterburner power setting. The drag model was created with CFD, where the mesh was built in ICEM Mesh and the simulations were run with the CFD solver M-Edge. Five Mach numbers in the range from 0.6 to 1.6 were simulated at an altitude of 12 km. The results showed that the afterbody drag generally decreased when increasing jet pressure ratio at both subsonic and supersonic velocities. At subsonic conditions, increasing nozzle area ratio for underexpanded nozzles would decrease the drag. Increasing nozzle area ratio for fully expanded or overexpanded nozzles would instead increase the drag to an intermediate point from where it would decrease. At supersonic condition, increasing nozzle area ratio would generally cause reduction in drag for all cases. The optimization showed that a net thrust increase of 0.02% to 0.09% could be gained for subsonic conditions while the supersonic optimization had negligible gain in thrust.

abstract: The objective of this study is to understand how to integrate conical spike external compression inlets with high bypass turbofan engines for application on future supersonic airliners. Many performance problems arise when inlets are matched with engines as inlets come with a plethora of limitations and losses that greatly affect an engine’s ability to operate. These limitations and losses include drag due to inlet spillage, bleed ducts, and bypass doors, as well as the maximum and minimum values of mass flow ratio at each Mach number that define when an engine can no longer function. A collection of tools was developed that allow one to calculate the raw propulsion data of an engine, match the propulsion data with an inlet, calculate the aerodynamic data of an aircraft, and combine the propulsion and aerodynamic data to calculate the installed performance of the entire propulsion system. Several trade studies were performed that tested how changing specific design parameters of the engine affected propulsion performance. These engine trade studies proved that high bypass turbofan engines could be developed with external compression inlets and retain effective supersonic performance. Several engines of efficient fuel consumption and differing bypass ratios were developed through the engine trade studies and used with the aerodynamic data of the Concorde to test the aircraft performance of a supersonic airliner using these engines. It was found that none of the engines that were tested came close to matching the supersonic performance that the Concorde could achieve with its own turbojet engines. It is possible to speculate from the results several different reasons why these turbofan engines were unable to function effectively with the Concorde. These speculations show that more tests and trade studies need to be performed in order to determine if high bypass turbofan engines can be developed for effective usage with supersonic airliners in any possible way. / Dissertation/Thesis / Run file and text files from the propulsion simulations performed in NPSS. / Input and output file used in EDET to generate aerodynamic data of Concorde. / Five column propulsion data of tested engines after inlet matching. / Masters Thesis Aerospace Engineering 2018