An internal, 3-D, viscous numerical flow simulation was performed on the rectangular-to-circular transition, bifurcated, 2D high speed civil transport engine inlet at takeoff. The objective of the study was to obtain the causes of flow distortion at the inlet fan face. The inlet was modeled with the centerbody in the fully collapsed takeoff position. A single block, 14Ox40x40 Polar grid topology of a 1/4 symmetry volume of the inlet was used in the simulation. The calculation employed the well established, robust PARC3D CFD code, which uses the full three-dimensional, Reynolds averaged, Navier-Stokes equations in strong conservation form. The flow was considered to be turbulent over the entire flow region. The turbulence model incorporated into PARC3D is the algebraic Baldwin Lomax model. Limitations existed in the local region where the flow interacts with the nose cone due to the inherent limitations of the turbulence model. The results showed that the flow throughout the inlet was well behaved. The turbulent boundary layers were thin and stayed attached to the surfaces of the inlet throughout the entire flowfield. A high pressure recovery was observed at the fan face. Radial distortion at the fan face was caused by thin boundary layers on the nose cone and cowl surfaces. Circumferential distortion was caused by pressure gradients produced by the wake of the splitter plate, located just upstream of the fan. / Master of Science

Serpentine diffusers have become a common feature in modern aircraft as they allow for certain benefits that are impossible with a traditional linear configuration. With the benefits, however, come certain disadvantages, namely flow distortions that reduce engine efficiency and decrease engine surge stability margins. These distortions are now being researched comprehensively to determine solutions for mitigating the adverse effects associated with them. This study investigates how varying the freestream turbulence intensity of the flow entering a serpentine diffuser affects the distortion patterns that are produced by the diffuser. Experiments were performed with a model serpentine diffuser on the Annular Cascade Facility of the Air Force Research Laboratory at Wright-Patterson Air Force Base. Hot wire anemometry was used to measure inlet turbulence, while static pressure probes located axially along the upper and lower surface of the model diffuser and total pressure probes located across the aerodynamic interface plane (AIP) were used to measure the distortion patterns of the flow passing through the diffuser. Varying levels of inlet freestream turbulence, ranging from 0 to 4%, were generated using square and round bar turbulence screens in three distinct test configurations. Axial static pressure measurements indicate that increasing turbulence slightly affects flow separation development downstream of the second turn. This effect is also seen at the AIP where the total pressure recovery increases with increasing level of inlet turbulence in the region of flow separation at the upper surface. The total pressure recovery along the lower surface is also seen to be increased with higher inlet turbulence. However, total pressure recovery increase across the entire AIP is almost negligible. Overall, the inlet freestream turbulence has a minor effect on the distortion patterns caused by the serpentine diffuser when compared with proven active inlet flow control methods.

No description available.

A technique to infer circumferential total pressure distortion intensity found in serpentine inlet airflow was established using wall-pressure fluctuation measurements. This sensing technique was experimentally developed for aircraft with serpentine inlets in a symmetric, level flight condition. The turbulence carried by the secondary flow field that creates the non-uniform total pressure distribution at the compressor fan-face was discovered to be an excellent indicator of the distortion intensity. A basic understanding of the secondary flow field allowed for strategic sensor placement to provide a distortion estimate with a limited number of sensors. The microphone-based distortion estimator was validated through its strong correlation with experimentally determined circumferential total pressure distortion parameter intensities (DPCP). This non-intrusive DPCP estimation technique was then used as a DPCP observer in a distortion feedback control system. Lockheed Martin developed the flow control technique used in this control system, which consisted of jet-type vortex generators that injected secondary flow to counter the natural secondary flow inherent to the serpentine inlet. A proportional-integral-derivative (PID) based control system was designed that achieved a requested 66% reduction in DPCP (from a DPCP of 0.023 down to 0.007) in less than 1 second. This control system was also tested for its ability to maintain a DPCP level of 0.007 during a quick ramp-down and ramp-up engine throttling sequence, which served as a measure of system robustness. The control system allowed only a maximum peak DPCP of 0.009 during the engine ramp-up. The successful demonstrations of this automated distortion control system showed great potential for applying this distortion sensing scheme along with Lockheed Martin's flow control technique to military aircraft with serpentine inlets. A final objective of this research was to broaden the non-intrusive sensing capabilities in the serpentine inlet. It was desired to develop a sensing technique that could identify control efforts that optimized the overall inlet aerodynamic performance with regards to both circumferential distortion intensity DPCP and average pressure recovery PR. This research was conducted with a new serpentine inlet developed by Lockheed Martin having a lower length-to-diameter ratio and two flow control inputs. A cost function based on PR and DPCP was developed to predict the optimal flow control efforts at several Mach numbers. Two wall-mounted microphone signals were developed as non-intrusive inlet performance sensors in response to the two flow control inputs. These two microphone signals then replaced the PR and DPCP metrics in the original cost function, and the new non-intrusive-based cost function yielded extremely similar optimal control efforts. / Ph. D.

Majority of the rain producing monsoon systems in India form or intensify over the Bay of Bengal and move onto the land. We expect the air-sea interaction to be a crucial factor in the frequent genesis and intensification of monsoon systems over the Bay. Knowledge of air-sea fluxes is essential in determining the air-sea interactions. However, the Bay remains a poorly monitored ocean basin and the state of the near surface conditions during the monsoon months remains to be studied in detail. For example, we do not know yet which among the various flux formulae used in the General circulation models are appropriate over the Bay since there are no direct measurements of surface fluxes here during the peak monsoon months. The present thesis aims towards filing that gap. In this thesis fluxes were computed using the Bulk method, Inertial dissipation method and direct covariance method. The flux comparisons were reasonable during certain flow conditions which are clearly identified. When these conditions are not met the differences among the fluxes from these methods can be larger than the inherent uncertainties' in the methods. Stratification, flow distortion and averaging time are the key variables that give rise to the differences in the fluxes. It is found that there are significant differences in the surface flux estimates computed from different atmospheric General Circulation Model bulk parameterization schemes. In this thesis, the flow gradients are estimated by taking advantage of the natural pitch and roll motion of the ship. A attempt is made to gain insight into the flow distortion and its influence on the fluxes. In our analysis it is found that the displacement of the streamlines is an important component in quantifying flow distortion.

This master thesis is focused on design and aerodynamic analysis of subsonic turbine engine inlet in NACA duct configuration for unmanned aircraft. The first part of this paper is methodics for design considerations for NACA duct, which is based on theoretical analysis of this type of inlet. The acquired knowledge is used to design an inlet for the specified unmanned aircraft that is subject of CFD analysis. The impact of deflectors is considered in the evaluation and the solution is compared to the S-duct inlet. The proposed inlet with deflectors meets DC60 distortion criterion for all specified cases and the pressure losses requirements are met for lower velocities. Based on the results, the recommended application is for aircraft that flies in optimal design conditions for most of the mission.