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NASA NTRS1997

NASA Ames Laminar Flow Supersonic Wind Tunnel (LFSWT) Tests of a 10 deg Cone at Mach 1.6

This work is part of the ongoing qualification of the NASA Ames Laminar Flow Supersonic Wind Tunnel (LFSWT) as a low-disturbance (quiet) facility suitable for transition research. A 10 deg cone was tested over a range of unit Reynolds numbers (Re = 2.8 to 3.8 million per foot (9.2 to 12.5 million per meter)) and angles of incidence (O deg to 10 deg) at Mach 1.6. The location of boundary layer transition along the cone was measured primarily from surface temperature distributions, with oil flow interferometry and Schlieren flow visualization providing confirmation measurements. With the LFSWT in its normal quiet operating mode, no transition was detected on the cone in the test core, over the Reynolds number range tested at zero incidence and yaw. Increasing the pressure disturbance levels in the LFSWT test section by a factor of five caused transition onset on the cone within the test core, at zero incidence and yaw. When operating the LFSWT in its normal quiet mode, transition could only be detected in the test core when high angles of incidence (greater than 5 deg) for cones were set. Transition due to elevated pressure disturbances (Tollmien-Schlichting) and surface trips produced a skin temperature rise of order 4 F (2.2 C). Transition due to cross flows on the leeward side of the cone at incidence produced a smaller initial temperature rise of only order 2.5 F (1.4 C), which indicates a slower transition process. We can conclude that these cone tests add further proof that the LFSWT test core is normally low-disturbance (pressure fluctuations greater than 0.1%), as found by associated direct flow quality measurements discussed in this report. Furthermore, in a quiet test environment, the skin temperature rise is sensitive to the type of dominant instability causing transition. The testing of a cone in the LFSWT provides an excellent experiment for the development of advanced transition detection techniques.

Wolf, Stephen W. D.↗
NASA NTRS1997

Real Time Quantitative 3-D Imaging of Diffusion Flame Species

A low-gravity environment, in space or ground-based facilities such as drop towers, provides a unique setting for study of combustion mechanisms. Understanding the physical phenomena controlling the ignition and spread of flames in microgravity has importance for space safety as well as better characterization of dynamical and chemical combustion processes which are normally masked by buoyancy and other gravity-related effects. Even the use of so-called 'limiting cases' or the construction of 1-D or 2-D models and experiments fail to make the analysis of combustion simultaneously simple and accurate. Ideally, to bridge the gap between chemistry and fluid mechanics in microgravity combustion, species concentrations and temperature profiles are needed throughout the flame. However, restrictions associated with performing measurements in reduced gravity, especially size and weight considerations, have generally limited microgravity combustion studies to the capture of flame emissions on film or video laser Schlieren imaging and (intrusive) temperature measurements using thermocouples. Given the development of detailed theoretical models, more sophisticated studies are needed to provide the kind of quantitative data necessary to characterize the properties of microgravity combustion processes as well as provide accurate feedback to improve the predictive capabilities of the computational models. While there have been a myriad of fluid mechanical visualization studies in microgravity combustion, little experimental work has been completed to obtain reactant and product concentrations within a microgravity flame. This is largely due to the fact that traditional sampling methods (quenching microprobes using GC and/or mass spec analysis) are too heavy, slow, and cumbersome for microgravity experiments. Non-intrusive optical spectroscopic techniques have - up until now - also required excessively bulky, power hungry equipment. However, with the advent of near-IR diode lasers, the possibility now exists to obtain reactant and product concentrations and temperatures non-intrusively in microgravity combustion studies. Over the past ten years, Southwest Sciences has focused its research on the high sensitivity, quantitative detection of gas phase species using diode lasers. Our research approach combines three innovations in an experimental system resulting in a new capability for nonintrusive measurement of major combustion species. FM spectroscopy or high frequency Wavelength Modulation Spectroscopy (WMS) have recently been applied to sensitive absorption measurements at Southwest Sciences and in other laboratories using GaAlAs or InGaAsP diode lasers in the visible or near-infrared as well as lead-salt lasers in the mid-infrared spectral region. Because these lasers exhibit essentially no source noise at the high detection frequencies employed with this technique, the achievement of sensitivity approaching the detector shot noise limit is possible.

Kane, Daniel J.↗
NASA NTRS1997

Elucidation of Free Radical and Optogalvanic Spectroscopy Associated with Microgravity Combustion via Conventional and Novel Laser Platforms

Combustion studies under both normal gravity and microgravity conditions depend a great deal on the availability and quality of the diagnostic systems used for such investigations. Microgravity phenomena are specially susceptible to even small perturbations and therefore non-intrusive diagnostic techniques are of paramount importance for successful understanding of reduced-gravity combustion phenomena. Several non-intrusive diagnostic techniques are available for probing and delineating normal as well as reduced gravity combustion processes, such as Rayleigh scattering, Raman scattering, Mie scattering, velocimetry, interferometric and Schlieren techniques, emission and laser-induced fluorescence (LIF) spectroscopy. Our approach is to use the LIF technique as a non-intrusive diagnostic tool for the study of combustion-associated free radicals and use the concomitant optogalvanic transitions to accomplish precise calibration of the laser wavelengths used for recording the excitation spectra of transient molecular species. In attempting to perform spectroscopic measurements on chemical intermediates, we have used conventional laser sources as well as new and novel platforms employing rare-earth doped solid-state lasers. Conventional (commercially available) sources of tunable UV laser radiation are extremely cumbersome and energy-consuming devices that are not very suitable for either in-space or in-flight (or microgravity drop tower) experiments. Traditional LIF sources of tunable UV laser radiation involve in addition to a pump laser (usually a Nd:YAG laser with an attached frequency-doubling stage), a tunable dye laser. In turn, the dye laser has to be provided with a dye circulation system and a subsequent stage for frequency-doubling of the dye laser radiation, together with a servo-tuning system (termed the 'Autotracker') to follow the wavelength changes and also an optical system (called the 'Frequency Separator') for separation of the emanating visible and UV beams. In contrast to this approach, we have devised an alternate arrangement for recording LIF excitation spectra of free radicals (following appropriate precursor fragmentation) that utilizes a tunable rare-earth doped solid state laser system with direct UV pumping. We have designed a compact and portable tunable UV laser system incorporating features necessary for both in-space and in-flight spectroscopy experiments. For the purpose of LIF excitation, we have developed an all-solid-state tunable UV laser that employs direct pumping of the solid-state UV-active medium employing UV harmonics from a Nd:YAG laser. An optical scheme with counterpropagating photolysis and excitation beams focused by suitable lenses into a reaction vacuum chamber was employed.

Misra, Prabhakar↗
NASA NTRS1997

Quantitative Measurement of Oxygen in Microgravity Combustion

A low-gravity environment, in space or in ground-based facilities such as drop towers, provides a unique setting for studying combustion mechanisms. Understanding the physical phenomena controlling the ignition and spread of flames in microgravity has importance for space safety as well as for better characterization of dynamical and chemical combustion processes which are normally masked by buoyancy and other gravity-related effects. Due to restrictions associated with performing measurements in reduced gravity, diagnostic methods which have been applied to microgravity combustion studies have generally been limited to capture of flame emissions on film or video, laser Schlieren imaging and (intrusive) temperature measurements using thermocouples. Given the development of detailed theoretical models, more sophisticated diagnostic methods are needed to provide the kind of quantitative data necessary to characterize the properties of microgravity combustion processes as well as provide accurate feedback to improve the predictive capabilities of the models. When the demands of space flight are considered, the need for improved diagnostic systems which are rugged, compact, reliable, and operate at low power becomes apparent. The objective of this research is twofold. First, we want to develop a better understanding of the relative roles of diffusion and reaction of oxygen in microgravity combustion. As the primary oxidizer species, oxygen plays a major role in controlling the observed properties of flames, including flame front speed (in solid or liquid flames), extinguishment characteristics, flame size and flame temperature. The second objective is to develop better diagnostics based on diode laser absorption which can be of real value in both microgravity combustion research and as a sensor on-board Spacelab as either an air quality monitor or as part of a fire detection system. In our prior microgravity work, an eight line-of-sight fiber optic system measured water vapor mole fractions in the NASA Lewis 2.2-sec Drop Tower. In that system, the laser and all electronics resided at the top of the drop tower and was connected via a fiber optic cable to the rig, on which a 'pitch and catch' set of fiber collimating lenses were used to transmit the laser beam across a jet diffusion flame. This system required eight independent detection/demodulation units and had poor spatial resolution. This research builds on this earlier work, resulting in an improved capability for quantitative, nonintrusive measurement of major combustion species. A vertical cavity surface-emitting diode laser (VCSEL) and a continuous spatial scanning method permit the measurement of temporal and spatial profiles of the concentrations and temperatures of molecular oxygen. High detection sensitivity is achieved with wavelength modulation spectroscopy (WMS). One-g experiments are performed using a slot diffusion flame. Microgravity measurements on a solid fuel (cellulose sheet) system are planned for the NASA Lewis 2.2-second Drop Tower Facility.

Silver, Joel A.↗
NASA NTRS1997

Gravitational Influences on Flame Propagation Through Non-Uniform Premixed Gas Systems

We have built an apparatus for measuring flame spread rates through non-homogeneous fuel-air mixtures as a function of layer thickness and concentration. The layer thickness is adjusted by controlling the diffusion time above a fuel-saturated porous media, while the concentration is controlled by the fuel temperature. Normal gravity tests with methanol have so far explored largely the effect of temperature, as well as the effects of various aspects of the apparatus. Good agreement with previous research has been obtained. We have also demonstrated the ability of a rainbow schlieren system to quantitatively measure fuel vapor concentrations in the static case.

Miller, Fletcher J.↗
NASA NTRS1997

Pressure Wave Propagation in a Screech Cycle

The screech noise generation process from supersonic under expanded jets, issuing from a sonic nozzle pressure ratio of 2.4 and 3.3 (expanded Mach number, M(sub j) = 1.10 and 1.42), is investigated experimentally. Spark Schlieren visualization at different phases of the screech cycle are clearly shown. The rms pressure fluctuation at the screech frequency is measured in the near field region by a traversing microphone.

Source record↗
NASA NTRS1997

Qualitative Flow Visualization of a 110-N Hydrogen/Oxygen Laboratory Model Thruster

The flow field inside a 110 N gaseous hydrogen/oxygen thruster was investigated using an optically accessible, two-dimensional laboratory test model installed in a high altitude chamber. The injector for this study produced an oxidizer-rich core flow, which was designed to fully mix and react inside a fuel-film sleeve insert before emerging into the main chamber section, where a substantial fuel film cooling layer was added to protect the chamber wall. Techniques used to investigate the flow consisted of spontaneous Raman spectra measurements, visible emission imaging, ultraviolet hydroxyl spectroscopy, and high speed schlieren imaging. Experimental results indicate that the oxygen rich core flow continued to react while emerging from the fuel-film sleeve, suggesting incomplete mixing of the hydrogen in the oxygen rich core flow. Experiments also showed that the fuel film cooling protective layer retained its integrity throughout the straight section of the combustion chamber. In the converging portion of the chamber, however, a turbulent reaction zone near the wall destroyed the integrity of the film layer, a result which implies that a lower contraction angle may improve the fuel film cooling in the converging section and extend the hardware lifetime.

deGroot, Wim A.↗
NASA NTRS1998

Detailed Experimental Data for CFD Code Validation

Recent interest in low cost, reliable access to space has generated increased interest in advanced technology approaches to space transportation systems. A key to the success of such programs lies in the development of advanced propulsion systems capable of achieving the performance and operations goals required for the next generation of space vehicles. One extremely promising approach involves the combination of rocket and air-breathing engines into a rocket-based combined-cycle engine (RBCC). A key element of that engine is the rocket ejector that is utilized in the zero to Mach two operating regime. Studies of RBCC engine concepts are not new and studies dating back thirty years are well documented in the literature. However, studies focused on the rocket ejector mode of the RBCC cycle are lacking. The present investigation utilizes an integrated experimental and computation fluid dynamics (CFD) approach to examine critical rocket ejector performance issues. In particular, the development of a predictive methodology capable of performance prediction is a key objective in order to analyze thermal choking and its control, primary/secondary pressure matching considerations, and effects of nozzle expansion ratio. To achieve this objective, the present study emphasizes obtaining new data using advanced optical diagnostics such as Schlieren photography and Raman spectroscopy, and CFD techniques to investigate mixing in the rocket ejector mode. A new research facility for the study of the rocket ejector mode has been developed. In the last milestone report, the operational capabilities of this research facility were described. In this milestone report, the experimentally obtained measurements for CFD code validation are presented and discussed.

Santoro, Robert J.↗
NASA NTRS1998

Scanning Mode Sensor for Detection of Flow Inhomogeneities

A scanning mode sensor and method is provided for detection of flow inhomogeneities such as shock. The field of use of this invention is ground test control and engine control during supersonic flight. Prior art measuring techniques include interferometry. Schlieren, and shadowgraph techniques. These techniques. however, have problems with light dissipation. The present method and sensor utilizes a pencil beam of energy which is passed through a transparent aperture in a flow inlet in a time-sequential manner so as to alter the energy beam. The altered beam or its effects are processed and can be studied to reveal information about flow through the inlet which can in turn be used for engine control.

Adamovsky, Grigory↗
NASA NTRS1998

Laser Pencil Beam Based Techniques for Visualization and Analysis of Interfaces Between Media

Traditional optical methods that include interferometry, Schlieren, and shadowgraphy have been used successfully for visualization and evaluation of various media. Aerodynamics and hydrodynamics are major fields where these methods have been applied. However, these methods have such major drawbacks as a relatively low power density and suppression of the secondary order phenomena. A novel method introduced at NASA Lewis Research Center minimizes disadvantages of the 'classical' methods. The method involves a narrow pencil-like beam that penetrates a medium of interest. The paper describes the laser pencil beam flow visualization methods in detail. Various system configurations are presented. The paper also discusses interfaces between media in general terms and provides examples of interfaces.

Adamovsky, Grigory↗
NASA NTRS1995

Fuel-Air Mixing Effect on Nox Emissions for a Lean Premixed-Prevaporized Combustion System

The lean premixed-prevaporized (LPP) concept effectively meets low nitrogen oxides (NOx) emission requirements for combustors with the high inlet temperature and pressure typical of the High-Speed Civil Transport (HSCT). For the LPP system fuel-air mixture uniformity is probably the most important factor for low NOx emissions. Previous studies have suggested that the fuel-air mixture uniformity can be severely affected by changing the number and configuration of fuel injection points. Therefore, an experimental study was performed to determine how the number of fuel injection points and their arrangement affect NOx emissions from an LPP system. The NOx emissions were measured by a gas-sampling probe in a flame-tube rig at the following conditions: inlet temperature of 810 K (1000 F), rig pressure of 10 atm, reference velocity of 150 ft/s, and residence time near 0.005 s. Additionally, a focused Schlieren diagnostic technique coupled with a high speed camera was used to provide a qualitative description of the spatial flow field.

Lee, Chi-Ming↗
NASA NTRS1961

The Effect of Air Bleed on the Heat Transfer and Pressure Distribution on 30 Conical Flares at a Mach Number of 6.8

The effect of a boundary-layer bleed at the start of a 30 deg half-angle flare upon the shape of the flow boundaries, the pressure distribution on the flare, and the heat transfer to the flare was studied at a Mach number of 6.8. The forebody was an ogive cylinder. Test Reynolds numbers, based on forebody length, ranged from 1 x 10(exp 6) to 7.4 x 10(exp 6). Schlieren photographs showed the effect produced upon the flow boundaries by varying the dimensions of the bleed in both the radial and axial directions and by blunting the lip at the leading edge of the flare. The heat transfer and pressure distribution on the flare were correlated with the shape and nature of the flow boundaries.

Crawford, Davis H.↗
NASA NTRS1961

Laminar Heat-Transfer and Pressure Measurements at a Mach Number of 6 on Sharp and Blunt 15 deg Half-Angle Cones at Angles of Attack Up to 90 deg

Two circular conical configurations having 15 deg half-angles were tested in laminar boundary layer at a Mach number of 6 and angles of attack up to 90 deg. One cone had a sharp nose and a fineness ratio of 1.87 and the other had a spherically blunted nose with a bluntness ratio of 0.1428 and a fineness ratio of 1.66. Pressure measurements and schlieren pictures of the flow showed that near-conical flow existed up to an angle of attack of approximately 60 deg. At angles of attack above 70 deg high-pressure areas were present near the base and the bow shock wave was considerably curved. Comparison of the results with simply applied theories showed that on the stagnation line pressures may be predicted by Newtonian theory, and heat transfer by local yawed-cylinder theory based on the yaw angle of the windward generator and the local radius of the cone. Base effects increased the heat transfer in a region extending forward approximately 15 to 30 percent of the windward generator. Circumferential pressure distributions were higher than the corresponding Newtonian distribution and a better prediction was obtained by modifying the theory to match the pressure at 90 deg from the windward generator to that on the surface of the cone at an angle of attack of 0 deg. Circumferential heat-transfer distributions were predicted satisfactorily up to about 60 deg from the stagnation line by using Lees' heat-flux distribution based on the Newtonian pressure. The effects of nose bluntness at large angles of attack were very small in the region beyond two nose radii from the point of tangency.

Conti, Raul J.↗
NASA NTRS1961

Study of Flow Over Oscillating Airfoil Models at a Mach Number of 7.0 in Helium

A wind-tunnel study of unsteady flow at a Mach number of 7 in helium has been conducted on several sting-mounted wedge, double-wedge, and flat-plate airfoil models with three different leading-edge radii. The data were obtained by taking high-speed schlieren motion pictures of the decaying motion of the model as it was released from an initial deflection. The shock-wave position observed on the sharp-leading-edge models during the oscillation was compared with that obtained by use of unsteady flow theory as well as steady-state theory. Comparison of theoretical results indicated that no unsteady-flow effects exist over the range of reduced frequencies k, 0.007 less than equal than k less than or equal 0.030, studied experimentally. The experimental results confirmed this finding as no unsteady-flow effects were detected in this reduced-frequency range. Comparison of shock-wave positions measured for the blunt models with those calculated by steady-state methods indicated fair agreement.

Arman, Ali↗
NASA NTRS1961

Static Longitudinal Stability of a Rocket Vehicle Having a Rear-Facing Step Ahead of the Stabilizing Fins

Tests were conducted at Mach numbers of 3.96 and 4.65 in the Langley Unitary Plan wind tunnel to determine the static longitudinal stability characteristics of a fin-stabilized rocket-vehicle configuration which had a rearward facing step located upstream of the fins. Two fin sizes and planforms, a delta and a clipped delta, were tested. The angle of attack was varied from 6 deg to -6 deg and the Reynolds number based on model 6 length was about 10 x 10. The configuration with the larger fins (clipped delta) had a center of pressure slightly rearward of and an initial normal-force-curve slope slightly higher than that of the configuration with the smaller fins (delta) as would be expected. Calculations of the stability parameters gave a slightly lower initial slope of the normal-force curve than measured data, probably because of boundary-layer separation ahead of the step. The calculated center of pressure agreed well with the measured data. Measured and calculated increments in the initial slope of the normal-force curve and in the center of pressure, due to changing fins, were in excellent agreement indicating that separated flow downstream of the step did not influence flow over the fins. This result was consistent with data from schlieren photographs.

Keynton, Robert J.↗
NASA NTRS1959

Chordwise Pressure Distributions over Several NACA 16-Series Airfoils at Transonic Mach Numbers up to 1.25

A two-dimensional wind-tunnel investigation of the pressure distributions over several NACA 16-series airfoils with thicknesses of 4, 6, 9, and 12 percent of the chord and design lift coefficients of 0, 0.2, 1 and 0.5 has been conducted in the Langley airfoil test apparatus at transonic Mach numbers from 0.7 to 1.25. The tests ranged in Reynolds number from 2.4 x 10 (exp 6) to 2.8 x 10 (exp 6) and in angle of attack from -10 to 12 degrees. Chordwise pressure distributions and schlieren flow photographs are presented without analysis.

Ladson, Charles L.↗
NASA NTRS1998

Characterization of Metalorganic Chemical Vapor Deposition

A series of experimental and numerical investigations to develop a more complete understanding of the reactive fluid dynamics of chemical vapor deposition were conducted. In the experimental phases of the effort, a horizontal CVD reactor configuration was used for the growth of InP at UVA and for laser velocimetry measurements of the flow fields in the reactor at LaRC. This horizontal reactor configuration was developed for the growth of III-V semiconductors and has been used by our research group in the past to study the deposition of both GaAs and InP. While the ultimate resolution of many of the heat and mass transport issues will require access to a reduced-gravity environment, the series of groundbased research makes direct contributions to this area while attempting to answer the design questions for future experiments of how low must gravity be reduced and for how long must this gravity level be maintained to make the necessary measurements. It is hoped that the terrestrial experiments will be useful for the design of future microgravity experiments which likely will be designed to employ a core set of measurements for applications in the microgravity environment such as HOLOC, the Fluid Physics/Dynamics Facility, or the Schlieren photography, the Laser Imaging Velocimetry and the Laser Doppler Velocimetry instruments under development for the Advanced Fluids Experiment Module.

Jesser, W. A.↗
NASA NTRS1999

Experimental and Computational Investigation of a Translating-Throat Single-Expansion-Ramp Nozzle

An experimental and computational study was conducted on a high-speed, single-expansion-ramp nozzle (SERN) concept designed for efficient off-design performance. The translating-throat SERN concept adjusts the axial location of the throat to provide a variable expansion ratio and allow a more optimum jet exhaust expansion at various flight conditions in an effort to maximize nozzle performance. Three design points (throat locations) were investigated to simulate the operation of this concept at subsonic-transonic, low supersonic, and high supersonic flight conditions. The experimental study was conducted in the jet exit test facility at the Langley Research Center. Internal nozzle performance was obtained at nozzle pressure ratios (NPR's) up to 13 for six nozzles with design nozzle pressure ratios near 9, 42, and 102. Two expansion-ramp surfaces, one concave and one convex, were tested for each design point. Paint-oil flow and focusing schlieren flow visualization techniques were utilized to acquire additional flow data at selected NPR'S. The Navier-Stokes code, PAB3D, was used with a two-equation k-e turbulence model for the computational study. Nozzle performance characteristics were predicted at nozzle pressure ratios of 5, 9, and 13 for the concave ramp, low Mach number nozzle and at 10, 13, and 102 for the concave ramp, high Mach number nozzle.

Deere, Karen A.↗