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At least 775 records · Page 43

Assessment of a flow-through balance for hypersonic wind tunnel models with scramjet exhaust flow simulation

The purpose of this investigation were twofold: first, to determine whether accurate force and moment data could be obtained during hypersonic wind tunnel tests of a model with a scramjet exhaust flow simulation that uses a representative nonwatercooled, flow-through balance; second, to analyze temperature time histories on various parts of the balance to address thermal effects on force and moment data. The tests were conducted in the NASA Langley Research Center 20-Inch Mach 6 Wind Tunnel at free-stream Reynolds numbers ranging from 0.5 to 7.4 x 10(exp 6)/ft and nominal angles of attack of -3.5 deg, 0 deg, and 5 deg. The simulant exhaust gases were cold air, hot air, and a mixture of 50 percent Argon and 50 percent Freon by volume, which reached stagnation temperatures within the balance of 111, 214, and 283 F, respectively. All force and moment values were unaffected by the balance thermal response from exhaust gas simulation and external aerodynamic heating except for axial-force measurements, which were significantly affected by balance heating. This investigation showed that for this model at the conditions tested, a nonwatercooled, flow-through balance is not suitable for axial-force measurements during scramjet exhaust flow simulation tests at hypersonic speeds. In general, heated exhaust gas may produce unacceptable force and moment uncertainties when used with thermally sensitive balances.

Huebner, Lawrence D.↗

A Navier-Stokes solver for high speed equilibrium flows and application to blunt bodies

This paper presents a finite element method for the solution of Navier-Stokes equations with the assumption of thermodynamic and chemical equilibrium. The method employs an upwind finite element technique with an implicit time-marching scheme for the solution, and uses an adaptively generated unstructured triangular mesh with several layers of quadrilateral elements near solid walls. The complexity associated with the assumption that the flow is in equilibrium is treated consistently, and the inviscid flux Jacobian matrices are derived. Several problems involving inviscid and viscous hypersonic flow past blunt are solved. Results are compared with other numerical results and experimental data, and generally good agreement is observed.

Prabhu, Ramadas K.↗

Upwind Algorithm For Parabolized Navier-Stokes Equations

Supersonic flow about cone calculated accurately. Report presents theoretical basis of computer code solving parabolized Navier-Stokes equations of supersonic and hypersonic flow. For increased accuracy in resolution of details of strong aerodynamic shocks, code incorporates implicit, finite-volume, upwind numerical-integration scheme. Performs well in numerical simulations of flows around simple bodies.

Lawrence, Scott L.↗

Raman-shifting an ArF excimer laser to generate new lines for obtaining optical diagnostic based information in flow fields

The application of tunable excimer lasers in combustion and flow diagnostics is almost routine nowadays. The properties of this laser system that enable density and temperature measurements in supersonic and hypersonic flow fields to be conducted are its high power, high repetition rate, and high spectral brightness. The limitation imposed by this system on these measurements is the paucity of lines in the wavelength region, the vacuum-ultraviolet, where species of interest, such as OH, N2, O2, H2, H2O, CO, NO, etc., are susceptible to electronic excitation to high-lying states. To circumvent this problem one normally resorts to nonlinear optical techniques such as frequency conversion via stimulated Raman scattering (SRS), more commonly known as Raman shifting or Raman mixing, to extend these nonintrusive and nonperturbing techniques to the shorter wavelengths in the VUV region and, for that matter, to longer wavelengths in the infrared region, if the need arises. The theoretical basis of SRS and its application are well documented in the literature. In essence, the Raman shift is a consequence of the inelastic scattering of the incident radiation by the sample. Most of the scattered radiation from the molecules of the sample is unchanged in frequency. However, a small fraction of the incident radiation is changed in frequency. This shift is a result of the fact that some of the incident photons on colliding with the molecules of the sample give up some of their energy and emerge with a lower energy resulting in the lower-frequency Stokes radiation. Other incident photons may increase their energy by colliding with the vibrationally excited molecules of the medium and emerge as higher-frequency antistokes radiation. The generation of the latter is the main objective of this project. The process, however, depends on several factors, including the beam quality of the pump laser, the cross-section of the gaseous medium, the gas pressure, and the ambient temperature of the gas near the focal region. Furthermore, since the Raman shifting process is polarization sensitive, it is necessary to have all of the laser energy in a single polarization. These factors were taken into consideration in the execution of the project. The implementation of the Raman shift was accomplished by focusing the 193 nm output of an ArF excimer laser (Lamda-Physik LPX 150) into a 1-meter long high pressure recirculating Raman cell filled with H2 gas. The laser system was modified in order to improve the mode quality of the pump beam to enhance the Raman shifting. To accomplish this feat, a prism beam expander and grating on the oscillator discharge provided wavelength tuning over the excimer gain profile. Furthermore, a triple-pass configuration, as opposed to unstable resonator optics, was employed in the operation of the amplifier cavity so that when the oscillator output radiation, focused by a 51-cm focal length fused silica lens through a 50 micron pinhole (serving as a spatial filter) and recollimated with a 25-cm focal length lens, was fed into the amplifier, it was injection locked, thereby providing tunable radiation with relatively low divergence. The forward scattered radiation emanating from the impingement of the modified pump beam on the Raman cell was detected using an energy meter after the latter had been separated from it using a dispersing prism.

Koker, Edmond B.↗

Non-Intrusive Measurement Techniques for Flow Characterization of Hypersonic Wind Tunnels

This manuscript describes the wide variety of optical measurement techniques for characterizing the flow in hypersonic wind tunnels. The introduction briefly describes different types of hypersonic wind tunnels, why they are used, and typical freestream conditions including fluctuating quantities. Description of these conditions defines the challenge for measurement techniques which have varying degrees of accuracy and precision, and work only in certain temperature, density and/or speed regimes. The rest of the manuscript is broken up into sections, by measurement technique. Each technique is described and then several examples are provided. The concluding chapter compares and contrasts different aspects of the measurement techniques including accuracy, precision, spatial resolution and temporal resolution.

Danehy, P. M.↗

An Approach for Simulating Transport of Nonspherical, Sublimating Particles in Rarefied Gas Flows via the Direct Simulation Monte Carlo Method

This paper extends the particle Green’s function approach of Gallis et al. [Phys. Fluids 13, 3482–3492 (2001)] to compute the force and heat transfer on nonspherical, sublimating particles in rarefied gas flows. Green’s functions are derived for rods and plates, which are representative of ice particle shapes, and a numerical procedure is developed for computing the Green’s functions for a general discretized convex shape. Additionally, a sublimation model is developed to simulate the evolution of ice particles in rarefied gas flows. These models are implemented in a direct simulation Monte Carlo code and demonstrated for the case of ice-particle-laden hypersonic flow over a ramp.

71 CLASSICAL AND QUANTUM MECHANICS, GENERAL PHYSIC↗

A user guide for the EMTAC-MZ CFD code

The computer code (EMTAC-MZ) was applied to investigate the flow field over a variety of very complex three-dimensional (3-D) configurations across the Mach number range (subsonic, transonic, supersonic, and hypersonic flow). In the code, a finite volume, multizone implementation of high accuracy, total variation diminishing (TVD) formulation (based on Roe's scheme) is used to solve the unsteady Euler equations. In the supersonic regions of the flow, an infinitely large time step and a space-marching scheme is employed. A finite time step and a relaxation or 3-D approximate factorization method is used in subsonic flow regions. The multizone technique allows very complicated configurations to be modeled without geometry modifications, and can easily handle combined internal and external flow problems. An elliptic grid generation package is built into the EMTAC-MZ code. To generate the computational grid, only the surface geometry data are required. Results obtained for a variety of configurations, such as fighter-like configurations (F-14, AVSTOL), flow through inlet, multi-bodies (shuttle with external tank and SRBs), are reported and shown to be in good agreement with available experimental data.

Szema, Kuo-Yen↗

Reynolds number effects on hypersonic characteristics of a 0.010-scale model of the 139-B shuttle orbiter

Longitudinal and lateral-directional stability characteristics of the 139-B orbiter (model 32-0) were obtained in a continuous flow hypersonic tunnel at Mach 10.3. Tests were made at Reynolds numbers of 1.04 million and 2.17 million (based on body length) over an angle of attack range of 12 deg to 36 deg at sideslip angles of 0 deg and minus 5 deg. Data were obtained at three elevon/body flap settings.

Bernot, P. T.↗

Evaluation of compressible-flow Preston tube calibrations

An experimental and analytical study has been made of the accuracy of various Preston tube calibration equations to determine local skin friction in two-dimensional supersonic and low-hypersonic flows. Experimental Preston tube calibration data from the present and previous studies were used to evaluate the calibration equations. The maximum value of the calibration parameters of the present data is more than an order of magnitude larger than that previously obtained; thereby, the evaluation of the various calibration equations is facilitated. The Preston tube technique was found to be very inaccurate in the low range of the calibration parameters. Above this range, there was a steady increase in accuracy with increasing values of the calibration parameters. No critical maximum tube diameter was found even with tubes about twice as large as the theoretical maximum allowable diameter. Of the two forms of the calibration equation previously existing, the logarithmic laws gave more accurate results than the power laws over an extended range of the calibration parameters in supersonic adiabatic flow.

Allen, J. M.↗

Graphite Ablation and Thermal Response Simulation Under Arc-Jet Flow Conditions

The Two-dimensional Implicit Thermal Response and Ablation program, TITAN, was developed and integrated with a Navier-Stokes solver, GIANTS, for multidimensional ablation and shape change simulation of thermal protection systems in hypersonic flow environments. The governing equations in both codes are demoralized using the same finite-volume approximation with a general body-fitted coordinate system. Time-dependent solutions are achieved by an implicit time marching technique using Gauess-Siedel line relaxation with alternating sweeps. As the first part of a code validation study, this paper compares TITAN-GIANTS predictions with thermal response and recession data obtained from arc-jet tests recently conducted in the Interaction Heating Facility (IHF) at NASA Ames Research Center. The test models are graphite sphere-cones. Graphite was selected as a test material to minimize the uncertainties from material properties. Recession and thermal response data were obtained from two separate arc-jet test series. The first series was at a heat flux where graphite ablation is mainly due to sublimation, and the second series was at a relatively low heat flux where recession is the result of diffusion-controlled oxidation. Ablation and thermal response solutions for both sets of conditions, as calculated by TITAN-GIANTS, are presented and discussed in detail. Predicted shape change and temperature histories generally agree well with the data obtained from the arc-jet tests.

Chen, Y.-K.↗

Effects of flow separation and cove leakage on pressure and heat-transfer distributions along a wing-cove-elevon configuration at Mach 6.9

External and internal pressure and cold-wall heating-rate distributions were obtained in hypersonic flow on a full-scale heat-sink representation of the space shuttle orbiter wing-elevon-cove configuration in an effort to define effects of flow separation on cove aerothermal environment as a function of cove seal leak area, ramp angle, and free-stream unit Reynolds number. Average free-stream Mach number from all tests was 6.9; average total temperature from all tests was 3360 R; free-stream dynamic pressure ranged from about 2 to 9 psi; and wing angle of attack was 5 deg (flow compression). For transitional and turbulent flow separation, increasing cove leakage progressively increased heating rates in the cove. When ingested mass flow was sufficient to force large reductions in extent of separation, increasing cove leakage reduced heating rates in the cove to those for laminar attached flow. Cove heating-rate distributions calculated with a method that assumed laminar developing channel flow agreed with experimentally obtained distributions within root-mean-square differences that varied between 11 and 36 percent where cove walls were parallel for leak areas of 50 and 100 percent.

Deveikis, W. D.↗

Recent skin friction techniques for compressible flows

A brief review is given of developments over the last decade in techniques for the measurement of skin friction in compressible airflows. Emphasis is placed on mean measurements beneath turbulent boundary layers in the supersonic and hypersonic flow regimes. Recent improvements in existing techniques such as the skin friction balance, Preston tube, sublayer fence, and heat transfer analogy are discussed. New or potential techniques including optical and acoustic measurements are also treated. Finally, new results are presented on the adaptation of the recently-developed laser interferometer skin friction meter for use in high-speed wind tunnels.

Settles, G. S.↗

A Preliminary Assessment of Step Effects on the BOLT Geometry in Mach 6 Flow

An infrared thermography measurement technique has been installed in the Langley Aerothermodynamic Laboratory. The system is calibrated for use with commercially available thermoplastics and ceramic materials, and to the surface temperature ranges when exposed to the facility hypersonic flow. Herein an uncertainty quantification is performed, revealing 5% surface temperature and 20% heat transfer uncertainty bands to two-sigma confidence. As a first application of the test measurement technique, an investigation was conducted to explore the correlation between the height of forward facing (FF) and rearward facing (RF) two-dimensional steps, free stream Reynolds number, and the onset of boundary layer transition from laminar to turbulent in Mach 6 flow on a 50% scale Boundary Layer Transition (BOLT) flight test vehicle geometry. Computational predictions were generated to identify potential x/L locations of transition onset for various flow conditions with a smooth outer mold line geometry and are compared to experimental results. The BOLT Step Test was conducted in the NASA Langley Aerothermodynamic Laboratory 20-inch Mach 6 wind tunnel at Reynolds numbers between 1.4 and 11.3 million conditions for various step heights. Data were collected using a developmental infrared thermography test technique and reduced to heat transfer for comparison with CFD solutions. This is a preliminary assessment of the developmental infrared thermography test technique used to acquire and reduce the experimental data. insight into the ongoing comparative analysis of the effects of FF and RF steps on boundary layer transition for the BOLT geometry in Mach 6 flow.

BOLT↗

Low Reynolds number flow past a blunt axisymmetric body at angle of attack

Low Reynolds number flow of an ideal gas over a blunt axisymmetric body of large half-angle at small angles of attack is investigated, for the case of laminar hypersonic flow. Time-varying viscous shock layer equations describing the flowfield are obtained from the full Navier-Stokes system by keeping terms to second order in the inverse square root of Re in both viscous and inviscid regions; the equations are valid for moderate to high Re. Drag, skin friction, and heating rates were obtained at small (or zero) angles of attack. Conditions experienced by planetary entry probes during the high-altitude (early) legs of an atmospheric entry trajectory are pertinent to the problem.

Kumar, A.↗

Modern Advances in Ablative TPS

Topics covered include: Physics of Hypersonic Flow and TPS Considerations. Destinations, Missions and Requirements. State of the Art Thermal Protection Systems Capabilities. Modern Advances in Ablative TPS. Entry Systems Concepts. Flexible TPS for Hypersonic Inflatable Aerodynamic Decelerators. Conformal TPS for Rigid Aeroshell. 3-D Woven TPS for Extreme Entry Environment. Multi-functional Carbon Fabric for Mechanically Deployable.

Venkatapathy, Ethiraj↗

Hypersonic aerothermal characteristics of a manned low finenes ratio shuttle booster

An investigation of a winged booster model having canards and an ascent configuration comprised of the booster mounted in tandem with an orbiter model has been conducted at Mach 10.2 in the continuous flow hypersonic tunnel. Longitudinal and lateral directional force characteristics were obtained over angle of attack ranges of -12 deg to 60 deg for the booster and -11 deg to 11 deg for the ascent configuration. Interference heating effects on the booster using the phase-change coating technique were determined at 0 deg angle of attack. Some oil flow photographs of the isolated booster and orbiter and ascent configuration are also presented.

Bernot, P. T.↗

Tapping the Brake for Entry, Descent, and Landing

A matrix of simulations of hypersonic flow over blunt entry vehicles with steady and pulsing retropropulsion jets is presented. Retropropulsion in the supersonic domain is primarily designed to reduce vehicle velocity directly with thrust. Retropropulsion in the hypersonic domain may enable significant pressure recovery through unsteady, oblique shocks while providing a buffer of reactant gases with relatively low total temperature. Improved pressure recovery, a function of Mach number squared and oblique shock angle, could potentially serve to increase aerodynamic drag in this domain. Pulsing jets are studied to include an additional degree of freedom to search for resonances in an already unsteady flow domain with an objective to maximize the time-averaged drag coefficient. In this paradigm, small jets with minimal footprints of the nozzle exit on the vehicle forebody may be capable of delivering the requisite perturbations to the flow. Simulations are executed assuming inviscid, symmetric flow of a perfect gas to enable a rapid assessment of the parameter space (nozzle geometry, plenum conditions, jet pulse frequency). The pulsed-jet configuration produces moderately larger drag than the constant jet configuration but smaller drag than the jet-off case in this preliminary examination of a single design point. The fundamentals of a new algorithm for this challenging application with time dependent, interacting discontinuities using the feature detection capabilities of Walsh functions are introduced.

Gnoffo, Peter A.↗

Comparison of thin-film resistance heat-transfer gages with thin-skin transient calorimeter gages in conventional hypersonic wind tunnels

Thin film gages deposited at the stagnation region of small (8.1-mm-diameter) hemispheres and gages mounted flush with the surface of a sharp-leading-edge flat plate were tested in the Langley continuous-flow hypersonic tunnel and in the Langley hypersonic CF4 tunnel. Two substrate materials were tested, quartz and a machinable glass-ceramic. Small hemispheres were also tested utilizing the thin-skin transient calorimeter technique usually employed in conventional tunnels. One transient calorimeter model was a thin shell of stainless steel, and the other was a thin-skin insert of stainless steel mounted into a hemisphere fabricated from a machinable-glass-ceramic. Measured heat-transfer rates from the various hemispheres were compared with one another and with predicted rates. The results demonstrate the feasibility and advantages of using-film resistance heat-transfer gages in conventional hypersonic wind tunnels over a wide range of conditions.

Miller, C. G., III↗