Molecular-Based Optical Measurement Techniques for Transition and Turbulence in High-Speed Flow

High-speed laminar-to-turbulent transition and turbulence affect the control of flight vehicles, the heat transfer rate to a flight vehicle's surface, the material selected to protect such vehicles from high heating loads, the ultimate weight of a flight vehicle due to the presence of thermal protection systems, the efficiency of fuel-air mixing processes in high-speed combustion applications, etc. Gaining a fundamental understanding of the physical mechanisms involved in the transition process will lead to the development of predictive capabilities that can identify transition location and its impact on parameters like surface heating. Currently, there is no general theory that can completely describe the transition-to-turbulence process. However, transition research has led to the identification of the predominant pathways by which this process occurs. For a truly physics-based model of transition to be developed, the individual stages in the paths leading to the onset of fully turbulent flow must be well understood. This requires that each pathway be computationally modeled and experimentally characterized and validated. This may also lead to the discovery of new physical pathways. This document is intended to describe molecular based measurement techniques that have been developed, addressing the needs of the high-speed transition-to-turbulence and high-speed turbulence research fields. In particular, we focus on techniques that have either been used to study high speed transition and turbulence or techniques that show promise for studying these flows. This review is not exhaustive. In addition to the probe-based techniques described in the previous paragraph, several other classes of measurement techniques that are, or could be, used to study high speed transition and turbulence are excluded from this manuscript. For example, surface measurement techniques such as pressure and temperature paint, phosphor thermography, skin friction measurements and photogrammetry (for model attitude and deformation measurement) are excluded to limit the scope of this report. Other physical probes such as heat flux gauges, total temperature probes are also excluded. We further exclude measurement techniques that require particle seeding though particle based methods may still be useful in many high speed flow applications. This manuscript details some of the more widely used molecular-based measurement techniques for studying transition and turbulence: laser-induced fluorescence (LIF), Rayleigh and Raman Scattering and coherent anti-Stokes Raman scattering (CARS). These techniques are emphasized, in part, because of the prior experience of the authors. Additional molecular based techniques are described, albeit in less detail. Where possible, an effort is made to compare the relative advantages and disadvantages of the various measurement techniques, although these comparisons can be subjective views of the authors. Finally, the manuscript concludes by evaluating the different measurement techniques in view of the precision requirements described in this chapter. Additional requirements and considerations are discussed to assist with choosing an optical measurement technique for a given application

Two-Point Focused Laser Differential Interferometry Second-Mode Measurements at Mach 6

A two-point focused laser differential interferometer (FLDI) is used to make measurements of density fluctuations on a 7 degree half-angle cone in a Mach 6 flow. The system was first characterized in the laboratory using laser induced breakdown to provide a well defined density fluctuation. The speed of the shock wave generated by the breakdown is verified using simultaneous high-speed schlieren. The FLDI system is then installed at the NASA Langley 20-Inch Mach 6 air tunnel to make measurements in the boundary layer of the 7 degree half-angle cone model and in the tunnel freestream for a unit Reynolds number range of 3.0 to 8.22 x 10(exp 6)/ft. Second-mode packets are visible in the spectra, with peak frequencies increasing linearly and peak amplitude increasing as a function of unit Reynolds number. The two-point measurement allows for the calculation of the second-mode wavepacket speeds, which propagate between 88% and 92% of the freestream velocity of the tunnel for all Reynolds numbers. The FLDI measurements are compared to surface-mounted fast-response pressure transducer measurements, where second-mode frequencies and wavepacket speeds are in good agreement

Simultaneous Conventional and Plenoptic Background Oriented Schlieren Imaging

Plenoptic Background Oriented Schlieren (BOS) is an emerging schlieren technique that is capable of providing 3D qualitative and quantitative information about density gradients present in a wide range of fluid dynamics problems. In this work, the fundamental concepts of plenoptic BOS are reviewed before discussing an open-air experiment with a buoyant plume where both conventional BOS and plenoptic BOS measurements were acquired simultaneously. Both cameras had the same field-of-view for all experiments, and three different focal plane arrangements were explored: (1) the focal plane was set to the background positions and the plume varied between 11 different positions relative to this focal plane, (2) the focal plane was set to 635- millimeters in front of the background position, and (3) the nominal focal plane varied while the position of the plume remained fixed. Such discussion will provide insight on how the two techniques compare, and what additional work is required to better understand the results provided by these two imaging systems

NO PLIF Study of Hypersonic Transition Over a Discrete Hemispherical Roughness Element

Nitric oxide (NO) planar laser-induced fluorescence (PLIF) has been use to investigate the hypersonic flow over a flat plate with and without a 2-mm (0.08-in) radius hemispherical trip. In the absence of the trip, for all angles of attack and two different Reynolds numbers, the flow was observed to be laminar and mostly steady. Boundary layer thicknesses based on the observed PLIF intensity were measured and compared with a CFD computation, showing agreement. The PLIF boundary layer thickness remained constant while the NO flowrate was varied by a factor of 3, indicating non-perturbative seeding of NO. With the hemispherical trip in place, the flow was observed to be laminar but unsteady at the shallowest angle of attack and lowest Reynolds number and appeared vigorously turbulent at the steepest angle of attack and highest Reynolds number. Laminar corkscrew-shaped vortices oriented in the streamwise direction were frequently observed to transition the flow to more turbulent structures

Analysis of the Two-Level NO PLIF Model for Low-Temperature High-Speed Flow Applications

The current work compares experimentally obtained nitric oxide (NO) laser-induced fluorescence (LIF) spectra with the equivalent spectra obtained analytically. The experimental spectra are computed from captured images of fluorescence in a gas cell and from a laser sheet passing through the fuel-air mixing flowfield produced by a high-speed fuel injector. The fuel injector is a slender strut that is currently being studied as a part of the Enhanced Injection and Mixing Project (EIMP) at the NASA Langley Research Center. This injector is placed downstream of a Mach 6 facility nozzle, which simulates the high Mach number airflow at the entrance of a scramjet combustor, and injects helium, which is used as an inert substitute for hydrogen fuel. Experimental planar (P) LIF is obtained by using a UV laser to excite fluorescence from the NO molecules that are present in either a gas cell or the facility air used for the EIMP experiments. The experimental data are obtained for several segments of the NO fluorescence spectrum. The selected segments encompass LIF lines with rotational quantum numbers appropriate for low-to-moderate temperature flows similar to those corresponding to the nominal experimental flow conditions. The experimental LIF spectra are then evaluated from the data and compared with those obtained from the theoretical models. The theoretical spectra are obtained from LIFBASE and LINUS software, and from a simplified version of the two-level fluorescence model. The equivalent analytic PLIF images are also obtained by applying only the simplified model to the results of the Reynolds-averaged simulations (RAS) of the mixing flowfield. Good agreement between the experimental and theoretical results provides increased confidence in both the simplified LIF modeling and CFD simulations for further investigations of high-speed injector performance using this approach

Three Component Velocity and Acceleration Measurement Using FLEET

The femtosecond laser electronic excitation and tagging (FLEET) method has been used to measure three components of velocity and acceleration for the first time. A jet of pure N2 issuing into atmospheric pressure air was probed by the FLEET system. The femtosecond laser was focused down to a point to create a small measurement volume in the flow. The long-lived lifetime of this fluorescence was used to measure the location of the tagged particles at different times. Simultaneous images of the flow were taken from two orthogonal views using a mirror assembly and a single intensified CCD camera, allowing two components of velocity to be measured in each view. These different velocity components were combined to determine three orthogonal velocity components. The differences between subsequent velocity components could be used to measure the acceleration. Velocity accuracy and precision were roughly estimated to be +/-4 m/s and +/-10 m/s respectively. These errors were small compared to the approx. 100 m/s velocity of the subsonic jet studied

Comparison of MSL RCS Jet Computations With Flow Visualization and Velocimetry

Numerical predictions of the Mars Science Laboratory (MSL) reaction control system (RCS) jets interacting with a Mach 10 hypersonic flow are compared to experimental nitric oxide (NO) planar laser-induced fluorescence (PLIF) data. The steady Reynolds Averaged Navier Stokes (RANS) equations using the Baldwin-Barth one-equation turbulence model were solved using the OVERFLOW code. The experimental PLIF data used for comparison consists of qualitative two-dimensional visualization images, qualitative reconstructed three-dimensional flow structures, and quantitative two-dimensional distributions of streamwise velocity. Through modeling of the PLIF signal equation, computational flow images (CFI) were produced and directly compared to the qualitative PLIF data. Post processing of the experimental and simulation data enabled the jet trajectory to be extracted for a more quantitative comparison. The two-dimensional velocity fields were reconstructed through interpolation of a series of single-component velocity profiles. Each distribution of single-component velocity was obtained using molecular tagging velocimetry (MTV). After validating the numerical model, the numerical solution was further examined to gain insight into hypersonic jet-in-crossflow interaction. Future NO-PLIF experiments are proposed based on this analysis

Review of Fluorescence-Based Velocimetry Techniques to Study High-Speed Compressible Flows

This paper reviews five laser-induced fluorescence-based velocimetry techniques that have been used to study high-speed compressible flows at NASA Langley Research Center. The techniques discussed in this paper include nitric oxide (NO) molecular tagging velocimetry (MTV), nitrogen dioxide photodissociation (NO2-to-NO) MTV, and NO and atomic oxygen (O-atom) Doppler-shift-based velocimetry. Measurements of both single-component and two-component velocity have been performed using these techniques. This paper details the specific application and experiment for which each technique has been used, the facility in which the experiment was performed, the experimental setup, sample results, and a discussion of the lessons learned from each experiment

Hypersonic Boundary Layer Measurements with Variable Blowing Rates Using Molecular Tagging Velocimetry

Measurements of mean and instantaneous streamwise velocity profiles in a hypersonic boundary layer with variable rates of mass injection (blowing) of nitrogen dioxide (NO2) were obtained over a 10-degree half-angle wedge model. The NO2 was seeded into the flow from a slot located 29.4 mm downstream of the sharp leading edge. The top surface of the wedge was oriented at a 20 degree angle in the Mach 10 flow, yielding an edge Mach number of approximately 4.2. The streamwise velocity profiles and streamwise fluctuating velocity component profiles were obtained using a three-laser NO2->NO photolysis molecular tagging velocimetry method. Observed trends in the mean streamwise velocity profiles and profiles of the fluctuating component of streamwise velocity as functions of the blowing rate are described. An effort is made to distinguish between the effect of blowing rate and wall temperature on the measured profiles. An analysis of the mean velocity profiles for a constant blowing rate is presented to determine the uncertainty in the measurement for different probe laser delay settings. Measurements of streamwise velocity were made to within approximately 120 gm of the model surface. The streamwise spatial resolution in this experiment ranged from 0.6 mm to 2.6 mm. An improvement in the spatial precision of the measurement technique has been made, with spatial uncertainties reduced by about a factor of 2 compared to previous measurements. For the quiescent flow calibration measurements presented, uncertainties as low as 2 m/s are obtained at 95% confidence for long delay times (25 gs). For the velocity measurements obtained with the wind tunnel operating, average single-shot uncertainties of less than 44 m/s are obtained at 95% confidence with a probe laser delay setting of 1 gs. The measurements were performed in the 31-inch Mach 10 Air Tunnel at the NASA Langley Research Center

Spectroscopic Measurement Techniques for Aerospace Flows

The conditions that characterize aerospace flows are so varied, that a single diagnostic technique is not sufficient for its measurement. Fluid dynamists use knowledge of similarity to help categorize and focus on different flow conditions. For example, the Reynolds number represents the ratio of inertial to viscous forces in a flow. When the velocity scales, length scales, and gas density are large and the magnitude of the molecular viscosity is low, the Reynolds number becomes large. This corresponds to large scale vehicles (e.g Airbus A380), fast moving objects (e.g. artillery projectiles), vehicles in dense fluids (e.g. submarine in water), or flows with low dynamic viscosity (e.g. skydiver in air). In each of these cases, the inertial forces dominate viscous forces, and unsteady turbulent fluctuations in the flow variables are observed. In contrast, flows with small length scales (e.g. dispersion of micro-particles in a solid rocket nozzle), slow moving objects (e.g. micro aerial vehicles), flows with low density gases (e.g. atmospheric re-entry), or fluids with a large magnitude of viscosity (e.g. engine coolant flow), all have low Reynolds numbers. In these cases, viscous forces become very important and often the flows can be steady and laminar. The Mach number, which is the ratio of the velocity to the speed of sound in the medium, also helps to differentiate types of flows. At very low Mach numbers, acoustic waves travel much faster than the object, and the flow can be assumed to be incompressible (e.g. Cessna 172 aircraft). As the object speed approaches the speed of sound, the gas density can become variable (e.g. flow over wing of Learjet 85). When the object speed is higher than the speed of sound (Ma > 1), the presences of shock waves and other gas dynamic features can become important to the vehicle performance (e.g. SR-71 Blackbird). In the hypersonic flow regime (Ma > 5), large changes in temperature begin to affect flow properties, causing real-gas effects to occur (e.g. X-43 Scramjet). At even higher Mach numbers, chemistry and nonequilibrium effects come into play (e.g. Startdust re-entry capsule), further complicating the measurement. These limits can be predicted by calculating the ratio of chemical and thermal relaxation time to the flow time scales. Other non-dimensional numbers can be used to further differentiate types of aerospace flows