Study of shock wave boundary layer interaction phenonemon using color schlieren and snapshot Proper Orthogonal Decomposition

The flow structure of shockwave boundary layer interaction (SWBLI) has been studied using Rainbow Schlieren Deflectometry (RSD), Ensemble Averaging, Fast Fourier Transform (FFT), and snapshot Proper Orthogonal Decomposition (POD) techniques. The Mach number of the approach free-stream was Mach = 3. Shockwave was generated with a 12° wedge. The color schlieren pictures are used to determine the transverse ray deflections at each pixel of the pictures taken using a high speed camera. The interaction region structure is described statistically with the ensemble average and, root mean square deflections. FFT technique is used to determine the dominant frequencies at different regions of the flow field. Results indicate that low frequency oscillations dominate the flow field. The POD technique results complement the findings of the ensemble averaging technique and show that distinct regions contain most of the energy in the flow field. These distinct regions are located around the reflected shock, around the shock wave reaching into the approach boundary layer and around the separation region over the edge of the separation bubble. (Published By University of Alabama Libraries

Study of shock wave boundary layer interaction using modal decomposition

Snapshot Proper Orthogonal Decomposition (SPOD) and Dynamic Mode Decomposition (DMD) have been used to study the flow features of oblique S hockwave Boundary Layer Interaction ( SWBLI). Ra inbow Schlieren Deflectometry (RSD) has been used to capture the flow data at 30,000 frames per second. The experimental setup produces a free stream flow of Mach 3.1. A 12° isosceles triangular prism attached to the roof of the test section generates the required oblique shock in the test section. The RSD pictures were used for data analysis to determine the transverse ray deflections at each pixel of the pictures. The analysis focuses on obtaining dominant flow regions in the flow and the corresponding mo des at which they occur. This modal information is used to comment on the unsteadiness associated with the flo

Design and testing of a high data rate instantaneous laser doppler velocity probe

An Instantaneous Laser Doppler Velocimetry (ILDV) probe is designed, built, and tested. The probe is capable of measuring a single component velocity data at a rate as high as two megahertz. The probe can be employed in high-speed and unsteady flows, especially where high data capture rate is needed such as in shock tubes, high-speed wind tunnels, and pulse detonation engines. However, the probe use is not restricted to only high speed flows. The probe, as designed, requires the flow direction to be known as it cannot discern the flow direction. Light scattered by particles illuminated by a laser beam in the flow is collected via backscattering. This light is transmitted to a breadboard housed in a container designed to insulate the system from sound, light, and vibration. The transmitted light is collimated and passed through a Michelson interferometer. Doppler frequency information contained in the beam is converted to light intensity variation using polarization optics. Two photomultiplier tubes are used to detect the light intensity variations. The output of the PM tubes are acquired using a high-speed data acquisition board, and analyzed in a PC to determine the flow velocity. (Published By University of Alabama Libraries

Multivariable optimization of liquid rocket engines using Particle Swarm algorithms

Liquid rocket engines are highly reliable, controllable, and efficient compared to other conventional forms of rocket propulsion. As such, they have seen wide use in the space industry and have become the standard propulsion system for launch vehicles, orbit insertion, and orbital maneuvering. Though these systems are well understood, historical optimization techniques are often inadequate due to the highly non-linear nature of the engine performance problem. In this thesis, a Particle Swarm Optimization (PSO) variant was applied to maximize the specific impulse of a finite-area combustion chamber (FAC) equilibrium flow rocket performance model by controlling the engine's oxidizer-to-fuel ratio and de Laval nozzle expansion and contraction ratios. In addition to the PSO-controlled parameters, engine performance was calculated based on propellant chemistry, combustion chamber pressure, and ambient pressure, which are provided as inputs to the program. The performance code was validated by comparison with NASA's Chemical Equilibrium with Applications (CEA) and the commercially available Rocket Propulsion Analysis (RPA) tool. Similarly, the PSO algorithm was validated by comparison with brute-force optimization, which calculates all possible solutions and subsequently determines which is the optimum. Particle Swarm Optimization was shown to be an effective optimizer capable of quick and reliable convergence for complex functions of multiple non-linear variables. (Published By University of Alabama Libraries

Estimation of morphing airfoil shapes and aerodynamic loads using artificial hair sensors

An active area of research in adaptive structures focuses on the use of continuous wing shape changing methods as a means of replacing conventional discrete control surfaces and increasing aerodynamic efficiency. Although many shape-changing methods have been used since the beginning of heavier-than-air flight, the concept of performing camber actuation on a fully-deformable airfoil has not been widely applied. A fundamental problem of applying this concept to real-world scenarios is the fact that camber actuation is a continuous, time-dependent process. Therefore, if camber actuation is to be used in a closed-loop feedback system, one must be able to determine the instantaneous airfoil shape, as well as the aerodynamic loads, in real time. One approach is to utilize a new type of artificial hair sensors (AHS) developed at the Air Force Research Laboratory (AFRL) to determine the flow conditions surrounding deformable airfoils. In this study, AHS measurement data will be simulated by using the flow solver XFoil, with the assumption that perfect data with no noise can be collected from the AHS measurements. Such measurements will then be used in an artificial neural network (ANN) based process to approximate the instantaneous airfoil camber shape, lift coefficient, and moment coefficient at a given angle of attack. Additionally, an aerodynamic formulation based on the finite-state inflow theory has been developed to calculate the aerodynamic loads on thin airfoils with arbitrary camber deformations. Various aerodynamic properties approximated from the AHS/ANN system will be compared with the results of the finite-state inflow aerodynamic formulation in order to validate the approximation approach. (Published By University of Alabama Libraries

Ultra-High Speed Rainbow Schlieren Deflectometry for whole field acoustic measurements in supersonic jets

The goal of this research is to advance the field of noise measurement techniques to better understand the fundamental guiding principles of noise generation. This is accomplished in this study by demonstrating the capabilities of Ultra-High Speed Rainbow Schlieren Deflectometry (UHS-RSD) technique to visualize and quantify, in real-time, sound waves propagating from a supersonic cold air jet. Basic optical theory states that light rays passing through varying density transparent medium undergo deviations from their original paths because of refraction. An experimental setup was developed to direct parallel white light rays through a supersonic air jet. UHS-RSD technique employs aforementioned phenomena and enables mapping the light deflection angles, a measure of deviation of a light ray from its original path. Ray deflection angle mapping process is realized through variation in color (hue) between a schlieren image with and without test medium. Since all information in the field of view can be acquired in real time, this technique provides a means to determine full field of view scalar properties of any transparent flow. The current experiment captures sound waves emanating from a supersonic cold air jet at high spatial and temporal resolution, while still maintaining the hue sensitivity needed to detect the small pressure fluctuations characteristic of sound waves. The sound probe data showing general maximum sound generation could be employed to support the visual UHS-RSD data where pressure gradient waves are seen propagating from the same jet noise source. Initial analysis reveals the UHS-RSD system to capture the acoustic field properties matching previous studies. Mach wave fronts are generated in small packets of 2 to 10 waves with small intermittency window. This research has shown that the UHS-RSD technique has the capability of capturing acoustic waves emanating from supersonic jets. A detailed discussion is presented on how the RSD system has been optimized to significantly improve sensitivity and the signal to noise ratio. (Published By University of Alabama Libraries

Computational analysis of diffuser performance for the Subsonic Aerodynamic Research Laboratory wind tunnel

The Air Force has expressed interest in improving the efficiency of the Subsonic Aerodynamic Research Laboratory (SARL) wind tunnel. In a previous analysis of losses throughout the tunnel, it was found that approximately thirty percent of pressure losses through the tunnel occurred at the exit of the tunnel (Britcher, 2011). The use of alternative diffuser geometries in reducing pressure losses at the exit of the tunnel and the computation of their efficiency improvement with respect to the original tunnel geometry and with respect to each other for the SARL wind tunnel are the focus of this research. Three different diffuser geometries were evaluated numerically using both the SolidWorks Flow Simulation add-on, and ANSYS FLUENT. For each of these geometries, a scaled down model was manufactured to be used for experimental validation in future work. Both the full size and small scale numerical models were evaluated with an inlet velocity of sixty meters per second. As the nature of the flow at this point in the wind tunnel is not known, both a uniform and fully developed turbulent flow profiles were evaluated for each design, both for the small scale models and the full size models, to determine pressure losses with respect to the varying flow types entering the diffusers. This research seeks to determine the effects of these different geometries on the flow downstream of the exit, and the possible energy savings associated with each design. In addition, it seeks to compare the numerical results obtained from both SolidWorks Flow Simulation and ANSYS FLUENT. (Published By University of Alabama Libraries

Experimental study on the effects of nose geometry on drag over axisymmetric bodies in supersonic flow

A new nose shape that was determined using the penetration mechanics to have the least penetration drag has been tested in the supersonic wind tunnel of the University of Alabama to determine the aerodynamic characteristics of this nose shape. The aerodynamic drag measured on the new nose shape and on four additional nose shapes are compared to each other. The results show that the new nose shape has the least aerodynamic drag. The measurements were made at Mach numbers ranging from 1.85 to 3.1. This study also required the maintenance of several components of the University of Alabama's 6-inch by 6-inch supersonic wind tunnel and modification of the existing data acquisition programs. These repairs and modifications included the repair and recalibration of the supersonic wind tunnel, repair of the four component force balance, and the modification of the tunnel's control program. (Published By University of Alabama Libraries