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An investigation was carried out to study the process of rotating stall inception in a low hub-tip ratio fan. Such fans are expected, based on an elementary analysis, to stall from the root. However, experimental evidence had led to the belief that the fans stalled from the tip. The effects of streamtube contraction were first studied and this was followed by an experimental investigation on an isolated rotor, with successive build modifications to increase the likelihood of rotating stall inception occurring at the root. A computer based streamline curvature method was used to study the effects of streamtube contraction and streamtube diffusion that commonly occur when a fan is operated at flows below its' design flow rate. The results indicated a reduced expectation for the root to stall first when compared to a simple 2-D flow analysis. Experimental measurements were then carried out to determine how the experimental local characteristics differed from the predicted characteristics. It was apparent that real fluid effects tended to steepen the root characteristic, thus enhancing the stability of the root. The tip characteristics tended to droop and become less stable. The enhancement of the root stability was also seen in the profiles of deviation angle. The axial Velocity contours at the rotor exit supported the conclusion that the root stability enhancement was caused by "centrifuging". To determine the actual radial location of rotating stall inception, an array of hot wires was used to record events during the inception transient. Inception was first detectable at the tip. This tip stalling behaviour persisted for all the build modifications. Measurements of unsteady pressure were also made to study the movement of the overall operating point since it was felt that this could continuously alternate between a pair of closely spaced characteristics. The results indicated that the fan operated along a unique characteristic. The overall conclusion was that a low hub-tip ratio fan shows a strong reluctance to stall at the root due to "centrifuging" of the blade boundary layer. The inception process appears to be dominated by events in the tip region.

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ABSTRACT DESIGN OF A CONNECTED PIPE TEST FACILITY FOR RAMJET APPLICATIONS SARISIN, Mustafa Nevzat M.S., Department of Mechanical Engineering Supervisor: Asst. Prof. Dr. Abdullah ULAS Co-Supervisor: Prof. Dr. Kahraman ALBAYRAK April 2005, 164 pages Development of the combustor of a ramjet can be achieved by connected pipe testing. Connected pipe testing is selected for combustor testing because pressure, temperature, Mach number, air mass flow rate can be simulated by this type of testing. Real time trajectory conditions and transition from rocket motor (booster) to ramjet operation can also be tested. The biggest advantage of connected pipe testing is the low operation cost and simplicity. Air mass flow rate requirement is less than the others which requires less air storage space and some components like supersonic nozzle and ejector system is not necessary. In this thesis, design of a connected pipe test facility is implemented. Three main systems are analyzed / air storage system, air heater system and test stand. Design of air storage system includes the design of pressure vessel and pressure &amp / flow regulation system. Pressure and flow regulation system is needed to obtain the actual flow properties that the combustor is exposed to during missile flight. Alternatives for pressure and air mass flow rate regulation are considered in this study. Air storage system designed in this thesis is 27.8 m3 at 50 bar which allows a test duration of 200 seconds at an average mass flow rate of 3 kg/s. Air heater system is utilized to heat the air to simulate the aerodynamic heating of the inlet. Several different combustion chamber configurations with different flame holding mechanisms are studied. The most efficient configuration is selected for this study. Combustion analysis of the air heater is performed by FLUENT CFD Code. Combustion process and air heater designs are validated using experimental data. Designed air heater system is capable of supplying air at a temperature range of 400-1000 K and mass flow rate range of 1.5-8 kg/s at Mach numbers between 0.1-0.5 and pressure between 2-8 bar. Finally the design of the test stand and ramjet combustor analysis are completed. 3D CAD models of the test stand are generated. Ramjet combustor that will be tested in the test setup is modeled and combustion analysis is performed by FLUENT CFD Code. The ramjet engine cruise altitude is 16 km and cruise Mach number is 3.5. Key-words: Air Breathing Engines, Ramjet, Connected Pipe, Direct Connect, Vitiator.

A new suboptimal guidance law design approach for aerospace vehicles is proposed in this thesis, followed by an advanced control design for supersonic air-breathing engines. The guidance law is designed using the newly developed Generalized Model Predictive Static Programming (G-MPSP), which is based on the continuous time nonlinear optimal control framework. The key feature of this technique is one-time backward propagation of a small-dimensional weighting matrix dynamics, which is used to update the entire control history. This key feature, as well as the fact that it leads to a static optimization problem, lead to its computational efficiency. It has also been shown that the existing model predictive static programming (MPSP), which is based on the discrete time framework, is a special case of G-MPSP. The G-MPSP technique is further extended to incorporate ‘input inequality constraints’ in a limited sense using the penalty function philosophy. Next, this technique has been developed also further in a ‘flexible final time’ framework to converge rapidly to meet very stringent final conditions with limited number of iterations. Using the G-MPSP technique in a flexible final time and input inequality constrained formulation, a suboptimal guidance law for a solid motor propelled carrier launch vehicle is successfully designed for a hypersonic mission. This guidance law assures very stringent final conditions at the injection point at the end of the guidance phase for successful beginning of the hypersonic vehicle operation. It also ensures that the angle of attack and structural load bounds are not violated throughout the trajectory. A second-order autopilot has been incorporated in the simulation studies to mimic the effect of the inner-loops on the guidance performance. Simulation studies with perturbations in the thrust-time behaviour, drag coefficient and mass demonstrate that the proposed guidance can meet the stringent requirements of the hypersonic mission. The G-MPSP technique in a fixed final time and input inequality constrained formulation has also been used for optimal guidance of an aerospace vehicle propelled by supersonic air-breathing engine, where the resulting thrust can be manipulated by managing the fuel flow and nozzle area (which is not possible in solid motors). However, operation of supersonic air-breathing engines is quite complex as the thrust produced by the engine is a result of very complex nonlinear combustion dynamics inside the engine. Hence, to generate the desired thrust, accounting for a fairly detailed engine model, a dynamic inversion based nonlinear state feedback control design has been carried out. The objective of this controller is to ensure that the engine dynamically produces the thrust that tracks the commanded value of thrust generated from the guidance loop as closely as possible by regulating the fuel flow rate. Simultaneously, by manipulating throat area of the nozzle, it also manages the shock wave location in the intake for maximum pressure recovery with sufficient margin for robustness. To filter out the sensor and process noises and to estimate the states for making the control design operate based on output feedback, an extended Kalman filter (EKF) based state estimation design has also been carried out and the controller has been made to operate based on estimated states. Moreover, independent control designs have also been carried out for the actuators so that their response can be faster. In addition, this control design becomes more challenging to satisfy the imposed practical constraints like fuel-air ratio and peak combustion temperature limits. Simulation results clearly indicate that the proposed design is quite successful in assuring the desired performance of the air-breathing engine throughout the flight trajectory, i.e., both during the climb and cruise phases, while assuring adequate pressure margin for shock wave management.