An experimental investigation is carried out to characterize the flow through a shallow single serpentine curved convergent nozzle with an aft deck under fully expanded and underexpanded conditions with nozzle pressure ratios ranging from 1.1 to 2.2 for cold flows by wall static pressure measurements and flow visualization of the exit jet. The study revealed initial deceleration of the flow due to the turning of the streamlines caused by the nozzle curvature. Asymmetric coefficient of pressure variation in the cross-stream direction across various nozzle pressure ratios was observed, which is caused by secondary flow created by wall curvature and transition. The acceleration of the core flow due to geometric transition, as well as the additional nozzle length, causes early transition to sonic and supersonic flows, resulting in the formation of shock cells at relatively lower Mach number in the absence of the aft deck, as can be seen from the schlieren images. However, the presence of the constant area extension section reduced the flow asymmetry. Furthermore, the presence of the aft deck prevented the formation of barrel shocks and is expected to prevent the formation Mach disk(s) at higher nozzle pressure ratios. Thus, the curvature, transition, and aft deck complemented each other.
Understanding the behavior of an aeroelastic system beyond the critical point is essential for effective implementation of any active control scheme since the control system design depends on the type of instability (bifurcation) the system encounters. Previous studies had found the aeroelastic system to enter into chaos beyond the point of instability. In the present work, an attempt has been made to carry out an experimental study on an aeroelastic model placed in a wind tunnel, to understand the behavior of aerodynamics around a wing section undergoing classical flutter. Wind speed was increased from zero until the model encountered flutter. Pressure at various locations along the surface of wing and acceleration at multiple points on the wing were measured in real time for the entire duration of experiment. A Leading Edge Separation Bubble (LSB) was observed beyond the critical point. The growing strength of the LSB with increasing wind speed was found to alter the aerodynamic moment acting on the system, which forced the system to enter into a second bifurcation. Based on the nature of the response, the system appears to undergo periodic doubling bifurcation rather than Hopf-bifurcation, resulting in chaotic motion. Eliminating the LSB can help in preventing the system from entering chaos. Any active flow control scheme that can avoid or counter the formation of leading edge separation bubble can be a potential solution to control the classical flutter.
Abstract Turbines remain one of the most efficient devices for extracting energy from a flowing fluid. In a gas turbine engine, axial flow turbines are used to extract energy from the working fluid and drive the compressor, to which they are mechanically connected. To maximize the performance of the axial flow turbine, it is necessary to carry out a design optimization of the components while suitably accounting for losses generated by secondary flows. An axial flow turbine rig is designed, fabricated, and installed to better understand and improve upon secondary flow models used in design procedures. The rig is driven by a blower operating at a constant speed, capable of delivering a maximum airflow rate of 0.4 kg/s and a maximum pressure rise of 500 mbar across the device. The axial flow turbine is mechanically connected to a dynamometer capable of operating at a full load capacity of 5 kW and a maximum rotational speed of 10,000 RPM. The axial flow turbine, housed between the blower and dynamometer, consists of nozzle guide vanes followed by a rotor. The design pressure ratio is chosen as 1.04, based on the blower delivery conditions and dynamometer specifications. For an initial design, a low-pressure ratio low rotor speed design was selected, allowing for easy installation and testing of the rotating components. The design space for the axial flow turbine was generated by varying flow and geometrical parameters in suitable steps, using a program written in MATLAB 2020a. Using the input variables and applying free vortex theory for three-dimensional blade design, the aerodynamic design of the axial flow turbine was carried out. The axial flow turbine design is experimentally tested with suitable pressure measurements at every station. Experiments are conducted for four different air mass flow rates. At each air mass flow, the rotor speed is varied by increasing/decreasing the dynamometer load. The data is recorded and compared with the design point. The difference between the design and measured performance parameters is observed to be within acceptable limits.
A periodic wave having a frequency that is an integral multiple of the fundamental power line frequency component is harmonics. They are the byproducts of modern electronics devices so it is necessary to mitigate the harmonics and offer techniques to mitigation of harmonics. This paper provides an explanation of the various harmonic mitigation techniques available to solve harmonic problems in three phase power systems. Included are the advantages and disadvantages of each method, their normal circuit connection as well as typical performance that can be expected when each method is properly employed.
Abstract During the static testing of a single spool turbojet engine (400 N Titan Engine) running on Jet A-1 as fuel, the starting sequence was unsuccessful, however, upon substituting the fuel with Diesel the engine started successfully. Till that point, the engine operation/performance was normal, and a series of tests were conducted for a total time of 5 hours. These 5 hours of testing spanned over five months during which ambient temperature varied from 5 °C to 37 °C, and relative humidity varied from 40% to 90%. The starting sequence was unsuccessful at the maximum ambient temperature (37 °C) and moderate relative humidity condition (50%). During the test, various parameters including pressure, temperature, rotor speed and mass flow rate were measured and recorded using a computer based data acquisition. For the failed start-up case, during the starting sequence, a sudden decrease in rotor speed was observed as the rotor reached ∼80% of the calibration value followed by flameout in combustion chamber. This sudden drop in rotor speed occurred due to mild-surge in compressor which continued up to flameout. The cause of mild-surge is attributed to the momentary blockage at the exit of the compressor caused by the combustion process. These pressure fluctuations propagated downstream leading to flame out. The combination of high ambient temperature (35 °C) and moderate humidity (∼50%) resulted in higher combustion chamber temperature which resulted in the momentary blockage at compressor exit during the transient operation. The substitution of Diesel as fuel resulted in lower combustion chamber temperature, and resolved the issue. It is important to note that, the mild-surge observed in this case is caused by downstream condition, and would occur only at some special situations (inlet conditions). Hence, it may be concluded, that the compressor surge in a jet engine may be caused by downstream components (under some special conditions), and design modifications to these downstream components may resolve the issue in those cases.
Prediction of wall pressure and heat flux due to aerodynamic heating in external flows is critical in the design of hypersonic vehicles as they are exposed to high aero-thermal loads. This paper examines and estimates the wall heat flux and pressure over a diamond airfoil in a supersonic flow (Mach = 5) for different angles of attack. We have obtained the wall heat flux results using the analytical correlations in order to justify the numerical results. Computed wall heat values agree with those given by correlations available in the literature. We find that the lower (high pressure) side of the airfoil experiences heat fluxes around 5 times the zero angles of attack values.
'%3e%3ccircle r='20' fill='%23008'/%3e%3ccircle r='17.5' fill='%23fff'/%3e%3ccircle r='3.5' fill='%23008'/%3e%3cg id='d'%3e%3cg id='c'%3e%3cg id='b'%3e%3cg id='a' fill='%23008'%3e%3ccircle r='.875' transform='rotate(7.5 -8.75 133.5)'/%3e%3cpath d='M0 17.5.6 7 0 2l-.6 5L0 17.5z'/%3e%3c/g%3e%3cuse xlink:href='%23a' transform='rotate(15)'/%3e%3c/g%3e%3cuse xlink:href='%23b' transform='rotate(30)'/%3e%3c/g%3e%3cuse xlink:href='%23c' transform='rotate(60)'/%3e%3c/g%3e%3cuse xlink:href='%23d' transform='rotate(120)'/%3e%3cuse xlink:href='%23d' transform='rotate(-120)'/%3e%3c/g%3e%3c/svg%3e)
