A low speed wind tunnel experimental investigation was conducted to determine the effectiveness of the leading edge pulsed blowing and the trailing edge jet blowing/ Gurney flap on the improvement of aerodynamic performance of an unmanned aerial vehicle at low Reynolds numbers. The wind tunnel tests for the leading edge pulsed jet blowing were conducted at 10%, 30% and 50% location of the chord length from the leading edge at a free stream velocity of 20 m/s. The jet momentum coefficient and the non-dimensional pulser frequency had been varied independently to investigate the effectiveness of the leading edge pulsed blowing. The trailing edge jet blowing tests were conducted at free stream velocity of 20 m/s at different jet momentum coefficients. The leading edge pulsed blowing showed a strong dependency of the actuator effectiveness on the jet momentum and the pulser frequency. The leading edge pulsed blowing had delayed the flow separation over the airfoil from an angle of attack of 17° to 22° with a docile stall for jet emanating at 10% location of the chord length for a jet momentum coefficient of 0.0275. The pulsed blowing at 50% chord location generated higher lift compared to the 10% location of the pulser with an abrupt stall at 19°. There was no evidence of the lift augmentation in the pre-stall angle of attack regime. The experimental results showed that the trailing edge jet flap was capable of generating significant roll moment at realistic jet momentum coefficients. The fluidic actuators were then integrated into the wings of a scale Extra 330 model airplane. The wind tunnel results for the leading edge pulsed blowing on the scale model indicated a delay in the stall of the airplane from an angle of attack of 12° to 21° with a 13% increase in the lift at take-off and landing speed of 17 m/s. The trailing edge jet actuators were also able to augment lift and demonstrate the roll control authority at low angle attacks at a cruising speed of 30 m/s.

A low speed wind tunnel experimental investigation was conducted to determine the effectiveness of the leading edge pulsed blowing and the trailing edge jet blowing/ Gurney flap on the improvement of aerodynamic performance of an unmanned aerial vehicle at low Reynolds numbers. The wind tunnel tests for the leading edge pulsed jet blowing were conducted at 10%, 30% and 50% location of the chord length from the leading edge at a free stream velocity of 20 m/s. The jet momentum coefficient and the non-dimensional pulser frequency had been varied independently to investigate the effectiveness of the leading edge pulsed blowing. The trailing edge jet blowing tests were conducted at free stream velocity of 20 m/s at different jet momentum coefficients. The leading edge pulsed blowing showed a strong dependency of the actuator effectiveness on the jet momentum and the pulser frequency. The leading edge pulsed blowing had delayed the flow separation over the airfoil from an angle of attack of 17° to 22° with a docile stall for jet emanating at 10% location of the chord length for a jet momentum coefficient of 0.0275. The pulsed blowing at 50% chord location generated higher lift compared to the 10% location of the pulser with an abrupt stall at 19°. There was no evidence of the lift augmentation in the pre-stall angle of attack regime. The experimental results showed that the trailing edge jet flap was capable of generating significant roll moment at realistic jet momentum coefficients. The fluidic actuators were then integrated into the wings of a scale Extra 330 model airplane. The wind tunnel results for the leading edge pulsed blowing on the scale model indicated a delay in the stall of the airplane from an angle of attack of 12° to 21° with a 13% increase in the lift at take-off and landing speed of 17 m/s. The trailing edge jet actuators were also able to augment lift and demonstrate the roll control authority at low angle attacks at a cruising speed of 30 m/s.

A large-scale model of a tiltrotor wing and flap was built and tested to determine how the size and spacing of fluidic actuators for active flow control should be scaled up from laboratory size models to aircraft size. Flow control was provided by sweeping jet actuators mounted in the flap follower. Chordwise pressure data was collected through pressure taps located along the model centerline. The maximum flap deflection to which the flow could be attached with actuation was compared to previous experiments on a 10% 3D model. An ideal actuator spacing was found and the momentum coefficient was comparable between the various sized models. These results were used in the design of an industrial wind tunnel-scale model of a wing built for cruise and tiltrotor-like download testing. These tests have shown that actuators located in the flap follower are effective in both the hover and cruise phases of flight.

A long aspect ratio synthetic jet is produced through an axial slit along part of the length of a cylinder. The jet is excited acoustically by a pair of loudspeakers mounted at the cylinder terminations. The study compares between the performance of two different slits with aspect ratios of 273 and 773. The comparison is based on the spanwise distribution of the mean jet velocity and phase between the jet velocity fluctuations and the excitation signal. Three different frequencies and amplitudes are used to excite the speakers covering the range of frequencies used in the control application. For both cases studied the mean centerline velocity of the jet increases with increasing the amplitude of the exciting signal, but decreases with increasing its frequency. Moreover, velocity deficits of up to 30% are evident as the midspan of the cylinder is approached from either end. Similar trends are also observed for the centerline phase distributions of the velocity fluctuations, with deficits of up to 130°. However, it is observed that for the long slit case the deficits in both the velocity and phase distributions are much larger than those for the short one. The synthetic jet is then mounted in the upstream cylinder of a tandem cylinder arrangement to be used as a control actuator for controlling the vibrations of the downstream cylinder. A simple feedback control mechanism is used at a Reynolds number of about 6.3x104. This Reynolds number corresponds to the case' where the downstream cylinder’s response is dominated with two frequency components, one at the resonance frequency of the cylinder, which is excited by broadband turbulence in the flow, and the other at the vortex shedding frequency. Both slits studied for the characterization experiments are used to compare their performance as control actuators. Both jets produce comparable reductions in the vibration of the downstream cylinder. A reduction of about 20% in the total RMS amplitude of the vibrations signal is achieved. This amounts to a reduction of about 50% in the resonant peak and an average value of about 40% in the vortex shedding peak. The optimal values of gain and time lag of the controller are then used to investigate the effect of the jet on the flow. It is found that the short slit jet produced an effect that was traced up to 1.875 diameters downstream, while the effect of the long slit jet dropped dramatically very close to the upstream cylinder. / Thesis / Master of Applied Science (MASc)

With increasing strain on the civil aviation industry to meet strict targets to reduce the adverse effects aviation has on the environment by 2050, significant advances in aircraft design and research are required. Aerodynamic improvements have been a focus for several decades now, however, current and future civil transport aircraft are based on traditional designs originating from the 1950s. Optimisation of aircraft external geometry for aerodynamic gain is reaching maturity and is becoming increasingly non-cost-effective. New advances in sensor and actuator technology has allowed for the development of active flow control (AFC) devices that have shown promising results in laboratory and even full-scale flight conditions, as seen by the joint NASA-Boeing ecoDemonstrator. One such device is the synthetic jet actuator (SJA), that synthesises periodic jets without the requirement for external air supply, while adding momentum to the surrounding flow. For this reason, SJAs are also referred to as zero-net-mass-flux actuators. There exists extensive work on the use of these devices for flow control applications in a laboratory setting. One of the key issues that remains unresolved, hindering successful aircraft application to-date, is the actuator self-noise generated. The noise level of SJAs can be so severe that they were rejected for application on the ecoDemonstrator in favour of a higher authority, quieter AFC device. SJAs were only considered for use in emergency situations on aircraft. Furthermore, the actuators were also not permitted to operate simultaneously at full power, which may severely limit scope for flow control on aircraft. Other applications that would benefit from SJAs include heat transfer for cooling in electronic devices. Studies in this field identify the same problem with noise levels of up to 73 dB reported. It is clear that work towards the self-noise reduction of SJAs is required to harness the full potential of this actuator technology. In the work presented, passive and active noise control measures in the form of lobed orifices and antiphase operation of two jets, respectively, on the noise reduction of SJAs are ii investigated. Noise sources of synthetic jet actuators include mechanical (diaphragm) and jet induced noise, where the focus of this work is on the latter type. Tests were conducted in quiescent conditions using jet velocity measurements, acoustic measurements, and flow visualisation. Tests were carried out using a single chamber SJA with variable cavity height and both circular and lobed orifices. These tests helped identify a SJA self-noise generation mechanism when using a circular orifice. This mechanism is characterised by a constant frequency behaviour visible in acoustic spectra for a specific jet Reynolds number range of 600 < Rej < 750 and Strouhal number range of 0.22 < St < 0.50. The geometries of the lobed orifices used in this work differ in lobe count and penetration. It was shown that a broadband noise reduction is possible with such orifices, with a maximum noise reduction of 14 dB at particular frequencies. The results indicate that a high number of lobes and penetration are preferred for noise reduction, however, at the expense of quickly dissipating downstream jet velocity. Flow visualisation reveals that this adverse effect is caused by enhanced mixing of lobed jets with ambient air that leads to earlier and more aggressive breakup of flow structures. A double chamber SJA is also used to demonstrate the noise attenuation through the antiphase operation of two cavities, caused by the interference pattern of the sound field of each source. The maximum reduction measured using this actuator configuration is 14 dB, depending on directivity.

Die vorliegende Arbeit widmet sich der, vorwiegend experimentellen, Untersuchung der Wirkung wandparalleler Lorentzkräfte in Strömungsrichtung auf Grenzschichtprofile und Körperumströmungen. Die Themen - Beeinflussung der Grenzschicht an ebenen Platten mit stationären Lorentzkräften - Kontrolle von Strömungsablösungen an Zylindern und symmetrischen Profilen mit stationären Lorentzkräften - Beeinflussung von Zylindernachläufen und abgelösten Tragflügelumströmungen mit zeitlich periodischen Lorentzkräften werden behandelt.

Die vorliegende Arbeit widmet sich der, vorwiegend experimentellen, Untersuchung der Wirkung wandparalleler Lorentzkräfte in Strömungsrichtung auf Grenzschichtprofile und Körperumströmungen. Die Themen - Beeinflussung der Grenzschicht an ebenen Platten mit stationären Lorentzkräften - Kontrolle von Strömungsablösungen an Zylindern und symmetrischen Profilen mit stationären Lorentzkräften - Beeinflussung von Zylindernachläufen und abgelösten Tragflügelumströmungen mit zeitlich periodischen Lorentzkräften werden behandelt.

The stable operating range of axial compressors is limited by the onset of rotating stall and surge. Mass injection upstream of the tip of an axial compressor rotor is a stability enhancement approach which can be effective in suppressing stall in tip-critical rotors, and thus increasing the operating range of compressors. In this article, investigations on active flow control related to the rotor tip gap sensitivity are discussed. The experiments were performed in a 1.5-stage low-speed research compressor. Measurements at part speed (80 per cent) and full speed (100 per cent) with varying injection rates are discussed. These tests were performed for two rotor tip clearances of 1.3 per cent and 4.3 per cent of rotor blade tip chord. Results on the compressor map, the flow field as well as transient measurements to identify the stall inception are discussed. Supplementary, the numerical results are compared to the experiments based on the configuration with the greatest benefit in operating range enhancement.

The research effort presented in this dissertation consists of employing active trailing edge blowing control to reduce the unsteady stator-rotor interaction in a turbofan simulator. Two active flow control systems with different wake sensing approaches are successfully implemented on the engine simulator. The first flow control system utilizes Pitot probes as flow sensors. Use of Pitot probes as sensors is appropriate as a first step toward a more in depth investigation of active trailing edge blowing control. An upper performance limit in terms of wake-filling can be obtained and serves as the baseline in evaluating other control systems with indirect wake sensors. The ability of the system to achieve effective wake filling when subjected to a change in inlet flow conditions demonstrates the feasibility and advantage of active flow control. Significant tonal noise reductions in the far field are also obtained. The second control system involves using microphones as indirect wake sensors. The significance of these acoustic sensing approaches is to provide a practical TEB approach for realistic engines implementations. Microphones are flush mounted on the inlet case to sense the tonal noise at the blade passing frequency. The first sensing approach only uses the tone magnitude while the second novel sensing approach utilizes both the tone magnitude and phase as error information. The convergence rate of the second sensing approach is comparable with that of the Pitot-probe based experiments. The acoustic results obtained from both sensing approaches agree well with those obtained using Pitot probes as sensors. In addition to the experimental part of this research, analytical studies are also conducted on the trailing edge blowing modeling using an aeroacoustic code. An analytical model for trailing edge blowing is first proposed. This model is then introduced into the two-dimensional aeroacoustic code to investigate effect of various trailing edge blowing managements in the tonal sound generation. / Ph. D.

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