Graduation Semester and Year
2019
Language
English
Document Type
Dissertation
Degree Name
Doctor of Philosophy in Aerospace Engineering
Department
Mechanical and Aerospace Engineering
First Advisor
Donald R. Wilson
Abstract
For a hypersonic flight mission, different flight regimes have been recognized. The successful and efficient operation of the aircraft to traverse through all the flight regimes requires the integration of various propulsion cycles into a single flow path. Using the phenomenon of detonation initiation and propagation, a multi-mode detonation based propulsion concept was proposed for hypersonic flight. Of the different modes proposed, it was recognized that the efficient operation of Normal Detonation Wave Engine (NDWE) mode and Oblique Detonation Wave Engine (ODWE) mode played an important role as they delivered thrust at critical parts of the trajectory. For this study, two different flight conditions representing the ODWE mode and NDWE mode are selected along a constant dynamic pressure trajectory of 47,880 N/m2. The ODWE mode is chosen as a design point and the flight Mach number chosen is 15 at an altitude of 42 km. The NDWE mode is the off-design condition and the flight Mach number representing this mode is 8.75 at an altitude of 34 km above sea level. An inviscid Euler simulation was carried out for the design point with an incoming combustion chamber Mach number of 6 which leads to the oblique detonation wave mode and the exit conditions at the expansion region are determined. The exit parameters of the expansion region are treated as inlet conditions into the nozzle. Nozzle inlet Mach number of 4.12 was determined from the simulation and using this Mach number, method of characteristics was used to design the nozzle contour for efficient expansion of the flow through the nozzle. Usually, hypersonic nozzles are large and they can be truncated at 40% of original length without significant loss of thrust. The designed nozzle is truncated at 40% of original length and CFD simulations are carried out to study the flow within the nozzle and also the flow interactions with the external flowfield. Using the mathematical model and geometry used for the design case, CFD simulations were carried out for off-design case with an incoming combustion chamber Mach number of 3.5. The CJ Mach number is greater than the incoming combustion chamber Mach number, leading to a moving detonation wave. The detonation wave movement is controlled and made to oscillate at a particular location downstream of the wedge by varying the stoichiometric ratio of the fuel-air mixture. The exit conditions at the expansion region are nearly constant because of this oscillation and the parameters at the exit are used as inlet conditions into the nozzle. Comparing the flow structures at the nozzle exhaust of the design and off-design conditions, similar shock structures are observed however, the plume appears to be longer and more voluminous in the design case when compared to the off-design case. Also, the plumes traverse a longer distance downstream of the nozzle before they mix into the external flowfield in case of the design case because of higher Mach number at the exit of the nozzle exhaust. This research sets the procedure to study the gas dynamic aspects of the flowfield from the combustion chamber all the way to the nozzle exhaust for a particular inlet combustion chamber conditions.
Keywords
Pulse detonation engine, Nozzle design, Hypersonic nozzle, Method of characteristics
Disciplines
Aerospace Engineering | Engineering | Mechanical Engineering
License
This work is licensed under a Creative Commons Attribution-NonCommercial-Share Alike 4.0 International License.
Recommended Citation
Kumar, Rahul, "Pulsed Detonation Engine Nozzle Design and Analysis" (2019). Mechanical and Aerospace Engineering Dissertations. 269.
https://mavmatrix.uta.edu/mechaerospace_dissertations/269
Comments
Degree granted by The University of Texas at Arlington