<div>To reduce the risk of using suction in a hypersonic quiet-tunnel nozzle design, this project tested micro-perforated suction sections to remove the boundary layer on an axisymmetric model in the Boeing/AFOSR Mach-6 Quiet Tunnel. The model was a cone-flare geometry tested at 0° angle of attack. The turn from the 7° half-angle cone to the flare was designed to prevent flow separation. The flare was designed to amplify the Görtler instability.</div><div><br></div><div>Five suction sections were designed with different perforation patterns and porosities. Four were successfully manufactured, but only the first of the four sections has been tested so far. The first suction section has pores drilled along straight lines with a nominal 5% porosity.</div><div><br></div><div>Measurements were made with temperature-sensitive paint and oil-flow visualization on a non-perforated blank to measure the baseline development of Görtler vortices on the flare. Although the signal-to-noise ratio of the measurement techniques were insufficient to measure the vortices, it was confirmed that the boundary layer is laminar for the entire model. Measurements with suction also did not show the Görtler vortices.</div><div><br></div><div>Surface pressure fluctuations were measured on the flare. Apparent second-mode waves were detected. The suction measurements showed a slight increase in second-mode peak frequency over the baseline results, as expected.</div><div><br></div><div>Concerns had been raised about acoustic noise that might be radiated from the suction section. Thus, fluctuations above the suction section were measured using a pitot probe and using focused-laser differential interferometry. The measurements during suction showed no noticeable increase in fluctuations compared to the baseline results.</div>

Transition to turbulence and flow control are studied by means of numerical simulations for different simple shear flows. Linear and non-linear optimisation methods using the Lagrange multiplier technique are employed. In the linear framework as objective function the standard disturbance kinetic energy is chosen and the constraints involve the linearised Navier–Stokes equations. We consider both the optimal initial condition leading to the largest disturbance energy growth at finite times and the optimal time-periodic forcing leading to the largest asymptotic response for the case of the flat plate boundary layer excluding the leading edge. The optimal disturbances for spanwise wavelengths of the order of the boundary layer thickness are streamwise vortices exploiting the lift-up mechanism to create streaks. For long spanwise wavelengths it is the Orr mechanism combined with the amplification of oblique wave packets that is responsible for the disturbance growth. Also linear optimal disturbances are computed around a leading edge and the effect of the geometry is considered. It is found that two-dimentional disturbances originating upstream, relative to the leading edge of the plate are inefficient at generating a viable disturbance, while three dimentional disturbances are more amplified. In the non-linear framework a new approach using ideas from non-equilibrium thermodynamics is developed. We determine the initial condition on the laminar/turbulent boundary closest to the laminar state. Starting from the general evolution criterion of non-equilibrium systems we propose a method to optimise the route to the statistically steady turbulent state, i.e. the state characterised by the largest entropy production. This is the first time information from the fully turbulent state is included in the optimisation procedure. The method is applied to plane Couette flow. We show that the optimal initial condition is localised in space for realistic flow domains, while the disturbance visits bent streaks before breakdown. Feedback control is applied to the bypass-transition scenario with high levels of free-stream turbulence. The flow is the flat-plate boundary layer. In this scenario low frequency perturbations enter the boundary layer and streamwise elongated disturbances emerge due to non-modal growth. The so-called streaky structures are growing in amplitude until they reach high enough energy levels and break down into turbulent spots via their secondary instability. When control is applied in the form of wall blowing and suction, the growth of the streaks is delayed, which implies a delay of the whole transition process. Additionally, a comparison with experimental work is performed demonstrating a remarkable agreement in the disturbance attenuation once the differences between the numerical and experimental setup are reduced. Open-loop control with wall travelling waves by means of blowing and suction is applied to a separating boundary layer. For downstream travelling waves we obtain a mitigation of the separation of the boundary layer while for upstream travelling waves a significant delay in the transition location accompanied by a modest reduction of the separated region. / QC 20110518

Boundary layer separation is an unwanted phenomenon in most technical applications, as for instance on airplane wings, ground vehicles and in internal flow systems. If separation occurs, it causes loss of lift, higher drag and energy losses. It is thus essential to develop methods to eliminate or delay separation.In the present experimental work streamwise vortices are introduced in turbulent boundary layers to transport higher momentum fluid towards the wall. This enables the boundary layer to stay attached at  larger pressure gradients. First the adverse pressure gradient (APG) separation bubbles that are to be eliminated are studied. It is shown that, independent of pressure gradient, the mean velocity defect profiles are self-similar when the scaling proposed by Zagarola and Smits is applied to the data. Then vortex pairs and arrays of vortices of different initial strength are studied in zero pressure gradient (ZPG). Vane-type vortex generators (VGs) are used to generate counter-rotating vortex pairs, and it is shown that the vortex core trajectories scale with the VG height h and the spanwise spacing of the blades. Also the streamwise evolution of the turbulent quantities scale with h. As the vortices are convected downstream they seem to move towards a equidistant state, where the distance from the vortex centres to the wall is half the spanwise distance between two vortices. Yawing the VGs up to 20° do not change the generated circulation of a VG pair. After the ZPG measurements, the VGs where applied in the APG mentioned above. It is shown that that the circulation needed to eliminate separation is nearly independent of the pressure gradient and that the streamwise position of the VG array relative to the separated region is not critical to the control effect. In a similar APG jet vortex generators (VGJs) are shown to as effective as the passive VGs. The ratio VR of jet velocity and test section inlet velocity is varied and a control effectiveness optimum is found for VR=5. At 40° yaw the VGJs have only lost approximately 20% of the control effect. For pulsed VGJs the pulsing frequency, the duty cycle and VR were varied. It was shown that to achieve maximum control effect the injected mass flow rate should be as large as possible, within an optimal range of jet VRs. For a given injected mass flow rate, the important parameter was shown to be the injection time t1. A non-dimensional injection time is defined as t1+ = t1Ujet/d, where d is the jet orifice diameter. Here, the optimal  t1+ was 100-200. / QC 20100825

An aluminum NACA 0018 airfoil testbed was constructed with 95 static pressure taps and 25 embedded microphones to enable novel time-resolved measurements of surface pressure. The main objective of this investigation is to utilize time-resolved surface pressure measurements to estimate salient flow characteristics in the separated flow region over the upper surface of an airfoil. The flow development over the airfoil was examined using hot wire anemometry and mean surface pressure for a range of Reynolds numbers from 80x103 to 200x103 and angles of attack from 0° to 18°. For these parameters, laminar boundary layer separation takes place on the upper surface and two flow regimes occur: (i) separation is followed by flow reattachment, so that a separation bubble forms and (ii) separation occurs without subsequent reattachment. Measurements of velocity and mean surface pressure were used to characterize the separated flow region and its effect on airfoil performance using the lift coefficient. In addition, the transition process and the evolution of disturbances were examined. The lift curve characteristics were found to be linked to the rate of change of the separation, transition, and reattachment locations with the angle of attack. For both flow regimes, transition was observed in the separated shear layer. Specifically, the amplification of disturbances within a band of frequencies in the separated shear layer resulted in laminar to turbulent transition. Validation of time-resolved surface pressure measurements was performed for Rec = 100x103 at α = 8° and α = 12°, corresponding to regimes of flow separation with and without reattachment, respectively. A comparative analysis of simultaneous velocity and time-resolved surface pressure measurements showed that the characteristics and development of velocity fluctuations associated with disturbances in the separated shear layer can be extracted from time-resolved surface pressure measurements. Specifically, within the separated flow region, the amplitude of periodic oscillations in the surface pressure signal associated with disturbances in the separated shear layer grew in the streamwise direction. In addition, the frequency at the spectral peak of the amplified disturbances in the separated shear layer was identified. Based on the results of the validation analysis, time-resolved surface pressure measurement analysis techniques were applied for a Reynolds number range from 60x103 to 130x103 and angles of attack from 6° to 16°. Within the separated flow region, the streamwise growth of surface pressure fluctuations is distinctly different depending on the flow regime. Specifically, within the separation bubble, the RMS surface pressure fluctuations increase in the streamwise direction and reach a peak just upstream of the reattachment location. The observed trend is in agreement with that observed for other separating-reattaching flows on geometries such as the forward and backward facing step and splitter plate with fence. In contrast to the separation bubble formation, when the separated shear layer fails to reattach to the airfoil surface, RMS surface pressure fluctuations increase in the streamwise direction with no maximum and the amplitude is significantly lower than those observed in the separation bubble. Surface pressure signals were further examined to identify the frequency, convective velocity, and spanwise uniformity of disturbances in the separated shear layer. Specifically, for both flow regimes, the fundamental frequency and corresponding Strouhal number exhibit a power-law dependency on the Reynolds number. Based on the available data for which velocity measurements were obtained in the separated flow region, the convective velocity matched the mean velocity at the wall-normal distance corresponding to the maximum turbulence intensity. A distinct increase in the convective velocity of disturbances in the separated shear layer was found when the airfoil was stalled in comparison to that found in the separation bubble. From statistical analysis of surface pressure signals in the spanwise direction, it was found that disturbances are strongly two-dimensional in the laminar portion of the separated shear layer and become three-dimensional through the transition process.

Aeroacoustic investigations regarding different types of helicopter noise have indicated that the most annoying noise is caused by impulsive blade surface pressure changes in descent or forward flight conditions. Blade Vortex Interaction (BVI) is one of the main phenomena producing significant impulsive noise by the unsteady fluctuation in blade loading due to the rapid change of induced velocity field during interaction with vortices shed from previous blades. The tip vortex core structure and the blade vortex miss distance were identified as having a primary influence on BVI. In this thesis, piezoelectrically modulated and/or vectored blowing at the rotor blade tip is theoretically investigated as an active technique for modifying the structure of the tip vortex core as well as for increasing blade vortex miss distance. The mechanisms of formation and convection of rotor blade tip vortices up to and beyond 360 degrees wake age are described based on the CFD results for the baseline cases of a hovering rotor with rounded and square tips. A methodology combining electromechanical and CFD modeling is developed and applied to the study of a piezoelectrically modulated and vectored blowing two-dimensional wing section. The thesis is focused on the CFD analysis of rotor flow with modulated tangential blowing over a rounded blade tip, and with steady mid-plane blade tip blowing, respectively. Computational results characterizing the far-wake flow indicate that for steady tangential blowing the miss distance can be doubled compared to the baseline case, which may lead to a significant reduction in BVI noise level if this trend shown in hover can be replicated in low speed forward flight. Moreover, near-wake flow analysis show that through modulated blowing a higher dissipation of vorticity can be obtained.

Turbulent mixing enhancement has received a great deal of attention in the fluid mechanics community in the last few decades. Generally speaking, mixing enhancement involves the increased dispersion of the fluid that makes up a flow. The current work focuses on mixing enhancement of an axisymmetric jet via high amplitude fluidic pulses applied at the nozzle exit with high aspect ratio actuator nozzles. The work consists of small scale clean jet experiments, small scale micro-turbine engine experiments, and full scale laboratory simulated core exhaust experiments using actuators designed to fit within the engine nacelle of a full scale aircraft. The small scale clean jet experiments show that mixing enhancement compared to the unforced case is likely due to a combination of mechanisms. The first mechanism is the growth of shear layer instabilities, similar to that which occurs with an acoustically excited jet except that, in this case, the forcing is highly nonlinear. The result of the instability is a frequency bucket with an optimal forcing frequency. The second mechanism is the generation of counter rotating vortex pairs similar to those generated by mechanical tabs. The penetration depth determines the extent to which this mechanism acts. The importance of this mechanism is therefore a function of the pulsing amplitude. The key mixing parameters were found to be the actuator to jet momentum ratio (amplitude) and the pulsing frequency, where the optimal frequency depends on the amplitude. The importance of phase, offset, duty cycle, and geometric configuration were also explored. The experiments on the jet engine and full scale simulated core nozzle demonstrated that pulse fluidic mixing enhancement was effective on realistic flows. The same parameters that were important for the cleaner small scale experiments were found to be important for the more realistic cases as well. This suggests that the same mixing mechanisms are at work. Additional work was done to optimize, in real time, mixing on the small jet engine using an evolution strategy.

The aerodynamic benefits of laminar flow have long made it a sought-after attribute in aircraft design. By laminarizing portions of an aircraft, such as the wing or empennage, significant reductions in drag could be achieved, reducing fuel burn rate and increasing range. In addition to environmental benefits, the economic implications of improved fuel efficiency could be substantial due to the upward trend of fuel prices. This is especially true for the commercial aviation industry, where fuel usage is high and fuel expense as a percent of total operating cost is high. Transition from laminar to turbulent flow can be caused by several different transition mechanisms, but the crossflow instability present in swept-wing boundary layers remains the primary obstacle to overcome. One promising technique that could be used to control the crossflow instability is the use of spanwise-periodic discrete roughness elements (DREs). The Flight Research Laboratory (FRL) at Texas A&M University has already shown that an array of DREs can successfully delay transition beyond its natural location in flight at chord Reynolds numbers of 8.0x10^6. The next step is to apply DRE technology at Reynolds numbers between 20x10^6 and 30x10^6, characteristic of transport aircraft. NASA's Environmentally Responsible Aviation Project has sponsored a transonic laminar-flow wing glove experiment further exploring the capabilities of DRE technology. The experiment will be carried out jointly by FRL, the NASA Langley Research Center, and the NASA Dryden Flight Research Center. Upon completion of a wing glove design, a thorough computational evaluation was necessary to determine if the design can meet the experimental requirements. First, representative CAD models of the testbed aircraft and wing glove were created. Next, a computational grid was generated employing these CAD models. Following this step, full-aircraft CFD flowfield calculations were completed at a variety of flight conditions. Finally, these flowfield data were used to perform boundary-layer stability calculations for the wing glove. Based on the results generated by flowfield and stability calculations, conclusions and recommendations regarding design effectiveness were made, providing guidance for the experiment as it moved beyond the design phase.

This study investigates the effect of chordwisely modulated tip injection on the flow and turbulence characteristics of the tip vortex through experimental measurements downstream of a rectangular half-wing that has an aspect ratio of three. This injection technique involves spanwise jets at the tip that are issued from a series of holes along the chord line normal to the freestream flow direction. The injection mass flow rate from each hole is individually controlled using computer driven solenoid valves and therefore the flow injection geometrical pattern at the tip can be adjusted to any desired waveform shape, with any proper injection velocity. The measurements are performed in a blow-down wind tunnel using Constant Temperature Anemometry and Kiel probe traverses as well as Stereoscopic Particle Image Velocimetry. Current data show consistent trends with v previously observed effects of steady uniform tip injection such as the upward and outward motion of the vortex as well as increased levels of turbulence within the vortex core. The vortex size gets bigger with injection and the total pressure levels get reduced significantly near the vortex core. The injection pattern also seems to affect the size of the wing wake as well as the wake entrainment characteristics of the tip vortex. Depending on the injection waveform pattern and injection momentum coefficient the helicoidal shape of the tip vortex also seems to get affected.

This study presents the results of an experimental study that investigates the effects of uniform/waveform tip injection along the camberline on the total pressure loss characteristics downstream of a row of Low Pressure Turbine (LPT) blades. The experiments are performed in a low speed cascade facility. This injection technique involves spanwise jets at the tip that are issued from a series of holes along the camber line normal to the freestream flow direction. The injection mass flow rate from each hole is individually controlled using computer driven solenoid valves and therefore the flow injection geometrical pattern at the tip can be adjusted to any desired waveform shape, and can be uniform as well as waveform along the camber. Measurements involve Kiel probe traverses for different injection scenarios 0.5 axial chords downstream of the blades. Results show that, instead of performing uniform mass injection along the camberline, by selecting an appropriate waveform injection pattern one can reduce the total loss levels of the blade, including the tip leakage loss as well as the wake losses.

This thesis presents the results of an experimental study that investigates the effects of steady vortex generator jets (VGJs) integrated to a low pressure turbine blade to control the laminar separation bubble occurring on the suction surface of the blade at low Reynolds numbers. The injection technique involves jets issued from the holes located near the suction peak of the test blade which is in the middle of a five-blade low speed linear cascade facility. Three injection cases are tested with different blowing ratio values ranging from low to high. Surface pressure and particle image velocimetry (PIV) measurements are performed. The results show that steady VGJ is effective in eliminating the laminar separation bubble. Also it is observed that to have fully developed attached boundary layer, blowing ratio should be chosen accordingly since a very thin separation zone still exists at low blowing ratios.