Active flow control (AFC) subscale experiments were conducted at the Lucas Wind Tunnel of the California Institute of Technology. Tests were performed on a generic vertical tail model at low speeds. Fluidic oscillators were used at the trailing edge of the main element (vertical stabilizer) to redirect the flow over the rudder and delay or prevent flow separation. Side force increases in excess of 50% were achieved with a 2% momentum coefficient ( ) input. The results indicated that a collective of about 1% could increase the side force by 30-50%. This result is achieved by reducing the spanwise flow on the swept back wings that contributes to early flow separation near their tips. These experiments provided the technical backdrop to test the full-scale Boeing 757 vertical tail model equipped with a fluidic oscillator system at the National Full-scale Aerodynamics Complex 40-by 80-foot Wind Tunnel, NASA Ames Research Center. The is shown to be an important parameter for scaling a fluidic oscillator AFC system from subscale to full-scale wind tunnel tests. The results of these tests provided the required rationale to use a fluidic oscillator AFC configuration for a follow-on flight test on the Boeing 757 ecoDemonstrator.

Femtosecond laser tagging is demonstrated for the first time in R134a (1,1,1,2-Tetrafluoroethane) gas, and in mixtures of R134a with small quantities of air. A systematic study of this tagging method is explored through the adjustment of gas pressure, mixture ratio and laser properties. It is found that the signal strength and lifetime are greatest at low pressures for excitation at both the 400 nm and 800 nm laser wavelengths. The relative intensities of two spectral peaks in the near-UV emission change as a function of gas pressure and can potentially be used for local pressure measurements. Single shot precision in pure R134a and R134a with 5% air is demonstrated in quiescent gas and at the exit of a subsonic pipe flow. One standard deviation (68%) of the uncertainty lies within 5 m/s of the mean velocity in a low pressure quiescent flow using a delay time of 3s, and 18 m/s in a 230 m/s flow using a delay of 5 s. The parameter space of these results are chosen to mimic conditions used in the NASA Langley Research Center's Transonic Dynamics Tunnel. The precision and signal lifetime demonstrate the feasibility of using this technique for measuring flowfields that induce airfoil flutter.

Reduced-order modeling using the Volterra series approach has been successfully applied in the past decades to weakly nonlinear aerodynamic and aeroelastic systems. However, aspects regarding the identification of the kernels associated with the convolution integrals of Volterra series can profoundly affect the quality of the resulting reduced-order model (ROM). An alternative method for their identification based on artificial neural networks is evaluated in this work. This relation between the Volterra kernels and the internal parameters of a time-delay neural network is explored for the application in the reduced-order modeling of nonlinear unsteady aerodynamic loads. An impulse-type Volterra-based ROM is also under consideration for comparison. All aerodynamic data used for the construction of the reduced-order models are obtained from computational fluid dynamics (CFD) simulations of the NACA 0012 airfoil using the Euler equations. Prescribed inputs in pitch and in plunge degrees of freedom at different free-stream Mach numbers are used to evaluate the range of applicability of the obtained models. For weakly nonlinear test cases, the modeling performance of the neural network Volterra ROM was comparable to the impulse-type ROM. Additional accuracy and adequate modeling of stronger nonlinearities, however, could only be attained with the inclusion of the neural network kernels of higher-order in the Volterra ROM. A generic expression is derived for the kernel function of -order from the internal parameters of a time-delay neural network.

Experimental measurements were performed on a swept flat-plate model with an airfoil leading edge and imposed chordwise pressure gradient to determine the effects of a backward-facing step on transition in a low-speed stationary crossflow-dominated boundary layer. Detailed hot-wire measurements were performed for three step heights ranging from 36 to 49% of the boundary-layer thickness at the step and corresponding to subcritical, nearly critical, and critical cases. In general, the step had a small localized effect on the growth of the stationary crossflow vortex, whereas the unsteady disturbance amplitudes increased with increasing step height. Intermittent spikes in instantaneous velocity began to appear for the two larger step heights. A physical explanation was provided for the mechanism leading to transition and the sudden movement in the transition front due to the critical steps. The large localized velocity spikes, which ultimately led to an intermittent breakdown of the boundary layer, were the result of nonlinear interactions of the different types of unsteady instabilities with each other and with the stationary crossflow vortices. Thus, the unsteady disturbances played the most important role in transition, but the stationary crossflow vortices also had a significant role via the modulation and the increased amplitude of the unsteady disturbances.

In this paper, we present a direct numerical simulation database of high-speed zero-pressure-gradient turbulent boundary layers developing spatially over a flat plate with nominal freestream Mach number ranging from 2.5 to 14 and wall-to-recovery temperature ranging from 0.18 to 1.0. The flow conditions of the DNS are representative of the operational conditions of the Purdue Mach 6 quiet tunnel, the Sandia Hypersonic Wind Tunnel at Mach 8, and the AEDC Hypervelocity Tunnel No. 9 at Mach 14. The DNS database is used to gauge the performance of compressibility transformations, including the classical Morkovin's scaling and strong Reynolds analogy as well as the newly proposed mean velocity and temperature scalings that explicitly account for wall heat flux. Several insights into the effect of direct compressibility are gained by inspecting the thermodynamic fluctuations and the Reynolds stress budget terms. Precomputed flow statistics, including Reynolds stresses and their budgets, will be available at the website of the NASA Langley Turbulence Modeling Resource, allowing other investigators to query any property of interest.

The "Reshotko-Tumin transition criterion" based on optimal transient growth successfully correlates laboratory measurements of roughness induced transition over blunt body configurations. Even though transient growth has not been conclusively linked to the measured onset of transition, the above correlation denotes the only available physics-based model for subcritical transition in blunt body flows, since the latter do not support any modal instabilities at typical experimental conditions. Unlike other established models based on empirical curve fits that are valid for a specific subclass of datasets, the optimal-growth-based transition criterion appears to provide a reasonable correlation with measurements in various wind tunnel and ballistic range facilities and for a broad range of surface temperature ratios. This paper is focused on optimal growth calculations that improve upon significant shortcomings of the computations underlying the Reshotko-Tumin correlation. The improved framework is applied to leeward transition over a spherical section forebody that was tested in the Mach 6 Adjustable Contour Expansion wind tunnel at Texas A&M University. The computed results highlight the significance of nonparallel basic state evolution, curvature terms, and an optimization procedure that varies both inflow and outflow locations of the transient growth interval. More important, the results indicate that the modified correlation is very close to its original form, and hence, that the accuracy of the transient-growth-based transition criterion is not compromised by using a more thorough theoretical framework. Yet the results also show that the optimal energy gain up to the predicted transition onset location can be rather small, highlighting the need to further investigate the optimal growth criterion for additional experimental configurations and to also uncover the in-depth physics underlying blunt body transition.

To study the azimuthal development of boundary-layer instabilities, a controlled, laser-generated perturbation was created in the freestream of the Boeing/U.S. Air Force Office of Scientific Research Mach 6 Quiet Tunnel. The freestream perturbation convected downstream in the wind tunnel to interact with a flared-cone model. The flared cone is a body of revolution bounded by a circular arc with a 3 m radius. Pressure transducers were used to measure a wave packet generated in the cone boundary layer by the freestream perturbation. Nine of these sensors formed three stations of azimuthal arrays and were used to determine the azimuthal variation of the wave packets in the boundary layer. The freestream laser-generated perturbation was positioned upstream of the model in three different configurations: along the centerline axis, offset from the centerline axis by 1.5 mm, and offset from the centerline axis by 3.0 mm. When the freestream perturbation was offset from the centerline of a flared cone with a 1.0 mm nose radius, a larger wave packet was generated on the side toward which the perturbation was offset. As a result, transition occurred earlier on that side. The offset perturbation did not have as large of an effect on the boundary layer of a nominally sharp flared cone.

A swept flat plate model with an imposed pressure gradient was experimentally investigated in a low-speed flow to determine the effect of a backward-facing step on transition in a stationary crossflow-dominated flow. Detailed hotwire measurements of boundary-layer flow were performed to investigate the upstream shift in transition due to a step height of 49% of the local unperturbed boundary-layer thickness. Increasing the initial stationary crossflow amplitude caused an upstream movement of the transition front for the backward-facing step case. The step caused a local increase in the growth of the stationary crossflow instabilities, but the stationary crossflow amplitude at transition was sufficiently low (<0.04 ) so that stationary crossflow was not solely responsible for transition. The unsteady velocity spectra downstream of the step were rich with unsteady disturbances in the 80- to 1500-Hz range. Three distinct families of disturbances were identified based on phase speed and wave angle, namely, a highly oblique disturbance (possibly traveling-crossflow-like), a Tollmien-Schlichting-wave-like disturbance, and a shear-layer instability. The stationary crossflow disturbances caused a modulation of the unsteady disturbances, resulting in spatially concentrated peaks in unsteady disturbance amplitude. This modulation of the unsteady disturbances is believed to be the reason for the upstream movement of the transition front with increasing stationary crossflow amplitude.

Supersonic boundary-layer receptivity to freestream acoustic disturbances is investigated by solving the Navier-Stokes equations for Mach 3.5 flow over a 7 deg half-angle cone. The freestream disturbances are generated from a wavy wall placed at the nozzle wall. The freestream acoustic disturbances radiated by the wavy wall are obtained by solving the linearized Euler equations. The results show that no noticeable instability modes are generated when the acoustic disturbances impinge the cone obliquely. The results show that the perturbations generated inside the boundary layer by the acoustic disturbances are the response of the boundary layer to the external forcing. The amplitude of the forced disturbances inside the boundary layer are about 2.5 times larger than the incoming field for zero azimuthal wave number, and they are about 1.5 times for large azimuthal wave numbers.

A parametric experimental study was performed with sweeping jet actuators (fluidic oscillators) to determine their effectiveness in controlling flow separation on an adverse pressure gradient ramp. Actuator parameters that were investigated include blowing coefficients, operation mode, pitch and spreading angles, streamwise location, and size. Surface pressure measurements and surface oilflow visualization were used to characterize the effects of these parameters on the actuator performance. 2D Particle Image Velocimetry measurements of the flow field over the ramp and hot-wire measurements of the actuator's jet flow were also obtained for selective cases. In addition, the sweeping jet actuators were compared to other well-known flow control techniques such as micro-vortex generators, steady blowing, and steady vortex-generating jets. The results confirm that the sweeping jet actuators are more effective than steady blowing and steady vortex-generating jets for this ramp configuration. The results also suggest that an actuator with a wider jet spreading (110 vs. 70 degrees) placed closer (2.3 vs. 7 boundary layer thickness upstream) to the flow separation location provides better performance. Different actuator sizes obtained by scaling down the actuator geometry produced different jet spreading. Scaling down the actuator (based on the throat dimensions) from 6.35 × 3.18 mm to 3.81 × 1.9 mm resulted in similar flow control performance; however, scaling down the actuator further to 1.9 × 0.95 mm reduced the actuator efficiency by reducing the jet spreading considerably. The results of this study provide insight that can be used to design and select the optimal sweeping jet actuator configuration for flow control applications.

Boundary-layer transition in hypersonic flows over a straight cone can be predicted using measured freestream spectra, receptivity, and threshold values for the wall pressure fluctuations at the transition onset points. Simulations are performed for hypersonic boundary-layer flows over a 7-degree half-angle straight cone with varying bluntness at a freestream Mach number of 10. The steady and the unsteady flow fields are obtained by solving the two-dimensional Navier-Stokes equations in axisymmetric coordinates using a 5-order accurate weighted essentially nonoscillatory (WENO) scheme for space discretization and using a third-order total-variation-diminishing (TVD) Runge-Kutta scheme for time integration. The calculated N-factors at the transition onset location increase gradually with increasing unit Reynolds numbers for flow over a sharp cone and remain almost the same for flow over a blunt cone. The receptivity coefficient increases slightly with increasing unit Reynolds numbers. They are on the order of 4 for a sharp cone and are on the order of 1 for a blunt cone. The location of transition onset predicted from the simulation including the freestream spectrum, receptivity, and the linear and the weakly nonlinear evolutions yields a solution close to the measured onset location for the sharp cone. The simulations overpredict transition onset by about twenty percent for the blunt cone.

The thrust of the dielectric barrier discharge plasma actuator is the plasma body force minus the wall shear force, and it equals the net induced momentum. Thrust measurement simplicity makes it a good metric of the aerodynamic performance for active flow control applications. Uncertainty and non-repeatability issues with conventional test setups motivated development of a novel suspended actuator test setup and a measurement methodology consisting of a burn-in procedure followed by frequency scans at constant voltages. This approach led to observation of negative values of thrust, or "antithrust," at low frequencies between 4 Hz and up to 64 Hz. The antithrust is proportional to the mean-squared voltage and is frequency independent. Departures from the parabolic antithrust curve are correlated with appearance of visible plasma discharges. The antithrust hypothesis is proposed. It states that the measured thrust is the sum of plasma thrust and antithrust. The magnitude of the antithrust depends on the actuator geometry, the materials, and the test installation. The dependence on test installation was validated by surrounding the actuator with a grounded large-diameter metal sleeve. A thrust data correction for antithrust enables meaningful comparisons between actuators at different installations. A strong dependence on humidity is also shown. The thrust significantly decreases with increasing humidity.

This paper presents mechanisms to explain, as well as mathematics to model, time-averaged spatially resolved amplitude observations of number density and number density unsteadiness in a Mach 10 flow as it transitions from the freestream, through a bow-shock wave, and into the gas cap created by a blunt-body model. The primary driver for bow-shock unsteadiness is freestream unsteadiness or "tunnel noise." Primary unsteadiness is bow-shock oscillation. It scales spatially with the number density first derivative and is modeled using a sech2() term. Secondary weaker unsteadiness begins as freestream unsteadiness and increases linearly in direct proportion to the gas number density across the bow shock and into the gas cap. This is the well-known amplification of the freestream turbulent kinetic energy mechanism and is modeled using a tanh() term. Total unsteadiness [fit using tanh() term + sech2() term] is expressed as the number density standard deviation and modeled as a linear combination of these two independent, simultaneous, and nonlinear unsteadiness mechanisms. Relationships between mechanism coefficients and various flowfield and wind-tunnel parameters are discussed. For example, bow-shock and gas cap oscillation amplitudes are linearly correlated with stagnation pressure and, by deduction, freestream unsteadiness.

Femtosecond laser electronic excitation tagging (FLEET) velocimetry was used to study the flowfield around a symmetric, transonic airfoil in the NASA Langley 0.3-m TCT facility. A nominal Mach number of 0.85 was investigated with a total pressure of 125 kPa and total temperature of 280 K. Two-components of velocity were measured along vertical profiles at different locations above, below, and aft of the airfoil at angles of attack of 0°, 3.5°, and 7°. Velocity profiles within the wake showed sufficient accuracy, precision, and sensitivity to resolve both the mean and fluctuating velocities and general flow physics such as shear layer growth. Evidence of flow separation is found at high angles of attack. Velocity measurements were assessed for their accuracy, precision, dynamic range, spatial resolution, and overall measurement uncertainty as they relate to the present experiments. Measurement precisions as low as 1 m/s were observed, while the velocity dynamic range was found to be nearly a factor of 500. The spatial resolution of between 1 mm and 5 mm was found to be primarily limited by the FLEET spot size and advection of the flow. Overall measurement uncertainties ranged from 3 to 4 percent.

This paper demonstrates a technique for locating the optimal control surface layout of an aeroservoelastic Common Research Model wingbox, in the context of maneuver load alleviation and active flutter suppression. The combinatorial actuator layout design is solved using ideas borrowed from topology optimization, where the effectiveness of a given control surface is tied to a layout design variable, which varies from zero (the actuator is removed) to one (the actuator is retained). These layout design variables are optimized concurrently with a large number of structural wingbox sizing variables and control surface actuation variables in order to minimize the sum of structural mass and actuator mass. The results demonstrate interdependencies between structural sizing patterns and optimal control surface layouts for both static and dynamic aeroelastic physics.

The performance of two wall models based on Reynolds-averaged Navier-Stokes is compared in large-eddy simulation of a high Reynolds number separating and reattaching flow over the NASA wall-mounted hump. Wall modeling significantly improves flow prediction on a coarse grid where the large-eddy simulation with the no-slip wall boundary condition fails. Low-order statistics from the wall-modeled large-eddy simulation are in good agreement with the experiment. Wall-pressure fluctuations from the resolved-scale solution are in good agreement with the experiment, whereas wall shear-stress fluctuations modeled entirely through the wall models appear to be significantly underpredicted. Although the two wall models produce comparable results in the upstream attached flow region, the nonequilibrium wall model outperforms the equilibrium wall model in the separation bubble and recovery region where the key assumptions in the equilibrium model are shown to be invalid.

The advancement of flow measurement techniques continues to extend experimental boundaries and thus significantly contributes to improving our understanding of both basic and applied aerodynamics. This is particularly apparent in the case of particle image velocimetry (PIV), where its application has furthered the existing knowledge in several areas of helicopter rotor aerodynamics. The complex nature of helicopter rotor flows presents unique challenges to experimentalists, including transonic flow, concentrated vortices and dynamic stall. To illustrate the impact of the technological advancements on the way helicopter aerodynamics is studied today, the development of PIV since the early nineties of the last century is reviewed and some recent PIV applications are described. Using examples of main rotor wakes, dynamic stall and flow control investigations, the capabilities of large-scale, time-resolved and volumetric PIV are summarized.

The accurate measurement of power consumption by dielectric barrier discharge plasma actuators is a challenge due to the characteristics of the actuator current signal. Microdischarges generate high-amplitude, high-frequency current spike transients superimposed on a low-amplitude, low-frequency current. A high-speed digital oscilloscope was used to measure the actuator power consumption using the shunt resistor method and the monitor capacitor method. The measurements were performed simultaneously and compared to each other in a time-accurate manner. It was found that low signal-to-noise ratios of the oscilloscopes used, in combination with the high dynamic range of the current spikes, make the shunt resistor method inaccurate. An innovative, nonlinear signal compression circuit was applied to the actuator current signal and yielded excellent agreement between the two methods. The paper describes the issues and challenges associated with performing accurate power measurements. It provides insights into the two methods including new insight into the Lissajous curve of the monitor capacitor method. Extension to a broad range of parameters and further development of the compression hardware will be performed in future work.

This paper describes a novel computational-fluid-dynamics-based numerical solution procedure for effective simulation of aerothermoacoustics problems with application to aerospace vehicles. A finite element idealization is employed for both fluid and structure domains, which fully accounts for thermal effects. The accuracies of both the fluid and structure capabilities are verified with flight- and ground-test data. A time integration of the structural equations of motion, with the governing flow equations, is conducted for the computation of the unsteady aerodynamic forces, which uses a transpiration boundary condition at the surface nodal points in lieu of the updating of the fluid mesh. Two example problems are presented herein to that effect. The first one relates to a cantilever wing with a NACA 0012 airfoil. The solution results demonstrate the effect of temperature loading that causes a significant increase in acoustic response. A second example, the hypersonic X-43 vehicle, is also analyzed; and relevant results are presented. The common finite element-based aerothermoelastic-acoustics simulation process, its applicability to the efficient and routine solution of complex practical problems, the employment of the effective transpiration boundary condition in the computational fluid dynamics solution, and the development and public domain distribution of an associated code are unique features of this paper.

In computational materials science, coarse-graining approaches often lack a priori uncertainty quantification (UQ) tools that estimate the accuracy of a reduced-order model before it is calibrated or deployed. This is especially the case in coarse-grained (CG) molecular dynamics (MD), where "bottom-up" methods need to run expensive atomistic simulations as part of the calibration process. As a result, scientists have been slow to adopt CG techniques in many settings because they do not know in advance whether the cost of developing the CG model is justified. To address this problem, we present an analytical method of coarse-graining rigid-body systems that yields corresponding intermolecular potentials with controllable levels of accuracy relative to their atomistic counterparts. Critically, this analysis: (i) provides a mathematical foundation for assessing the quality of a CG force field without running simulations; and (ii) provides a tool for understanding how atomistic systems can be viewed as appropriate limits of reduced-order models. Simulated results confirm the validity of this approach at the trajectory level and point to issues that must be addressed in coarse-graining fully non-rigid systems.