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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 487 records · Page 27
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NASA NTRS1992

Application of computational fluid dynamics and laminar flow technology for improved performance and sonic boom reduction

A discussion is given of the many factors that affect sonic booms with particular emphasis on the application and development of improved computational fluid dynamics (CFD) codes. The benefits that accrue from interference (induced) lift, distributing lift using canard configurations, the use of wings with dihedral or anhedral and hybrid laminar flow control for drag reduction are detailed. The application of the most advanced codes to a wider variety of configurations along with improved ray-tracing codes to arrive at more accurate and, hopefully, lower sonic booms is advocated. Finally, it is speculated that when all of the latest technology is applied to the design of a supersonic transport it will be found environmentally acceptable.

Bobbitt, Percy J.↗
NASA NTRS1992

Pitot survey of exhaust flow field of a 2-D scramjet nozzle at Mach 6 with air or freon and argon used for exhaust simulation

A pitot-rake survey of the simulated exhaust of a half-span scramjet nozzle model was conducted in the Langley 20-Inch Mach 6 Tunnel to provide an additional data set for computational fluid dynamics (CFD) code comparisons. A wind-tunnel model was tested with a 26-tube pitot rake that could be manually positioned along the mid-semispan plane of the model. The model configuration had an external expansion surface of 20 degrees and an internal cowl expansion of 12 degrees; tests were also performed with a flow fence. Tests were conducted at a free-stream Reynolds number of approximately 6.5 x 10(exp 6) per foot and a model angle of attack of -0.75 degrees. The two exhaust gas mediums that were tested were air and a Freon 12-argon mixture. Each medium was tested at two jet total pressures at approximately 28 and 14 psia. This document presents the flow-field survey results in graphical as well as tabular form, and several observations concerning the results are discussed. The surveys reveal the major expected flow-field characteristics for each test configuration. For a 50-percent freon 12 and 50-percent argon mixture by volume (Fr-Ar), the exhaust jet pressures were slightly higher than those for air. The addition of a flow fence slightly raised the pitot pressure for the Fr-Ar mixture, but it produced little change for air. For the Fr-Ar exhaust, the plume was larger and the region between the shock wave and plume was smaller.

Monta, William J.↗
NASA NTRS1992

Species and temperature measurement in H2/O2 rocket flow fields by means of Raman scattering diagnostics

Validation of Computational Fluid Dynamics (CFD) codes developed for prediction and evaluation of rocket performance is hampered by a lack of experimental data. Non-intrusive laser based diagnostics are needed to provide spatially and temporally resolved gas dynamic and fluid dynamic measurements. This paper reports the first non-intrusive temperature and species measurements in the plume of a 110 N gaseous hydrogen/oxygen thruster at and below ambient pressures, obtained with spontaneous Raman spectroscopy. Measurements at 10 mm downstream of the exit plane are compared with predictions from a numerical solution of the axisymmetric Navier-Stokes and species transport equations with chemical kinetics, which fully model the combustor-nozzle-plume flowfield. The experimentally determined oxygen number density at the centerline at 10 mm downstream of the exit plane is four times that predicted by the model. The experimental number density data fall between those numerically predicted for the exit and 10 mm downstream planes in both magnitude and radial gradient. The predicted temperature levels are within 10 to 15 percent of measured values. Some of the discrepancies between experimental data and predictions result from not modeling the three dimensional core flow injection mixing process, facility back pressure effects, and possible diffuser-thruster interactions.

Degroot, Wim A.↗
NASA NTRS1992

Overview of high performance aircraft propulsion research

The overall scope of the NASA Lewis High Performance Aircraft Propulsion Research Program is presented. High performance fighter aircraft of interest include supersonic flights with such capabilities as short take off and vertical landing (STOVL) and/or high maneuverability. The NASA Lewis effort involving STOVL propulsion systems is focused primarily on component-level experimental and analytical research. The high-maneuverability portion of this effort, called the High Alpha Technology Program (HATP), is part of a cooperative program among NASA's Lewis, Langley, Ames, and Dryden facilities. The overall objective of the NASA Inlet Experiments portion of the HATP, which NASA Lewis leads, is to develop and enhance inlet technology that will ensure high performance and stability of the propulsion system during aircraft maneuvers at high angles of attack. To accomplish this objective, both wind-tunnel and flight experiments are used to obtain steady-state and dynamic data, and computational fluid dynamics (CFD) codes are used for analyses. This overview of the High Performance Aircraft Propulsion Research Program includes a sampling of the results obtained thus far and plans for the future.

Biesiadny, Thomas J.↗
NASA NTRS1992

Comparison of methodologies for describing relaxation in nonequilibrium gaseous systems

The heat transfer process in hypervelocity vehicles is dominated by nonequilibrium gas dynamics. One model used in computational fluid dynamics (CFD) codes to predict hypervelocity heat transfer is the 'two-temperature' model. An analysis has been made to test the validity of the two-temperature model for predicting another nonequilibrium phenomenon, sound absorption and deviation of signal speed in a high temperature gas. It is found that the two temperature model's prediction capabilities degenerate with increasing temperature. These results are felt to have significance concerning the two-temperature's ability to predict heat transfer in hypervelocity flows.

Schreiber, Willard C.↗
NASA NTRS1993

Laser Rayleigh and Raman Diagnostics for Small Hydrogen/oxygen Rockets

Localized velocity, temperature, and species concentration measurements in rocket flow fields are needed to evaluate predictive computational fluid dynamics (CFD) codes and identify causes of poor rocket performance. Velocity, temperature, and total number density information have been successfully extracted from spectrally resolved Rayleigh scattering in the plume of small hydrogen/oxygen rockets. Light from a narrow band laser is scattered from the moving molecules with a Doppler shifted frequency. Two components of the velocity can be extracted by observing the scattered light from two directions. Thermal broadening of the scattered light provides a measure of the temperature, while the integrated scattering intensity is proportional to the number density. Spontaneous Raman scattering has been used to measure temperature and species concentration in similar plumes. Light from a dye laser is scattered by molecules in the rocket plume. Raman spectra scattered from major species are resolved by observing the inelastically scattered light with linear array mounted to a spectrometer. Temperature and oxygen concentrations have been extracted by fitting a model function to the measured Raman spectrum. Results of measurements on small rockets mounted inside a high altitude chamber using both diagnostic techniques are reported.

Degroot, Wilhelmus A.↗
NASA NTRS1992

Analysis of film cooling in rocket nozzles

Progress during the reporting period is summarized. Analysis of film cooling in rocket nozzles by computational fluid dynamics (CFD) computer codes is desirable for two reasons. First, it allows prediction of resulting flow fields within the rocket nozzle, in particular the interaction of the coolant boundary layer with the main flow. This facilitates evaluation of potential cooling configurations with regard to total thrust, etc., before construction and testing of any prototype. Secondly, CFD simulation of film cooling allows for assessment of the effectiveness of the proposed cooling in limiting nozzle wall temperature rises. This latter objective is the focus of the current work. The desired objective is to use the Finite Difference Navier Stokes (FDNS) code to predict wall heat fluxes or wall temperatures in rocket nozzles. As prior work has revealed that the FDNS code is deficient in the thermal modeling of boundary conditions, the first step is to correct these deficiencies in the FDNS code. Next, these changes must be tested against available data. Finally, the code will be used to model film cooling of a particular rocket nozzle. The third task of this research, using the modified code to compute the flow of hot gases through a nozzle, is described.

Woodbury, Keith A.↗
NASA NTRS1991

The lightcraft project

Rensselaer Polytechnic Institute has been developing a transatmospheric 'Lightcraft' technology which uses beamed laser energy to propel advanced shuttle craft to orbit. In the past several years, Rensselaer students have analyzed the unique combined-cycle Lightcraft engine, designed a small unmanned Lightcraft Technology Demonstrator, and conceptualized larger manned Lightcraft - to name just a few of the interrelated design projects. The 1990-91 class carried out preliminary and detailed design efforts for a one-person 'Mercury' Lightcraft, using computer-aided design and finite-element structural modeling techniques. In addition, they began construction of a 2.6 m-diameter, full-scale engineering prototype mockup. The mockup will be equipped with three robotic legs that 'kneel' for passenger entry and exit. More importantly, the articulated tripod gear is crucial for accurately pointing at, and tracking the laser relay mirrors, a maneuver that must be performed just prior to liftoff. Also accomplished were further design improvements on a 6-inch-diameter Lightcraft model (for testing in RPI's hypersonic tunnel), and new laser propulsion experiments. The resultant experimental data will be used to calibrate Computational Fluid Dynamic (CFD) codes and analytical laser propulsion models that can simulate vehicle/engine flight conditions along a transatmospheric boost trajectory. These efforts will enable the prediction of distributed aerodynamic and thruster loads over the entire full-scale spacecraft.

Messitt, Don G.↗
NASA NTRS1992

Nonintrusive spectroscopic techniques for supersonic/hypersonic aerodynamics and combustion diagnostics

This paper presents an overview of the primary nonintrusive diagnostic techniques being developed by the NASA Langley Research Center to address the validation needs of Computational Fluid Dynamic (CFD) codes. The techniques include absorption in the UV and IR, Laser Induced Fluorescence, electron beam fluorescence, and a number of scattering techniques including Rayleigh, spontaneous Raman, and several coherent Raman spectroscopies. Most of the techniques are highly specialized, require complex data interpretation, and can satisfy only a few of the CFD needs. For these reasons, the evolving trend in flowfield diagnostics appears to favor a mode in which the diagnostic researcher, the experimental aerodynamicist, and the CFD community jointly define experiments based on the aeronautical requirements and on available diagnostic techniques.

Exton, R. J.↗
NASA NTRS1993

The validation and application of numerical modelling to supersonic mixing and reacting flows

This study demonstrates that the computational fluid dynamics (CFD) code GASP can be used to predict the features for flowfields representative of scramjet combustors. The validation program undertaken was composed of two parts: a low enthalpy mixing study and a simplified autoignition study. The low enthalpy mixing study was done using low angle helium injection into a Mach 6 airstream over a flat plate. Both matched and overpressurized injection cases were numerically modelled. The agreement with experimental data for the farfield bulk mixing in both cases was reasonable. The combustion studies modelled premixed hydrogen-air reaction with the results being evaluated qualitatively. GASP proved to be highly sensitive to the chemical species present for flow near the ignition temperature, but provided reasonable qualitative results for a subsonic flameholder model.

Vitt, Paul H.↗
NASA NTRS1993

Axisymmetric computational fluid dynamics analysis of a film/dump-cooled rocket nozzle plume

Prediction of convective base heating rates for a new launch vehicle presents significant challenges to analysts concerned with base environments. The present effort seeks to augment classical base heating scaling techniques via a detailed investigation of the exhaust plume shear layer of a single H2/O2 Space Transportation Main Engine (STME). Use of fuel-rich turbine exhaust to cool the STME nozzle presented concerns regarding potential recirculation of these gases to the base region with attendant increase in the base heating rate. A pressure-based full Navier-Stokes computational fluid dynamics (CFD) code with finite rate chemistry is used to predict plumes for vehicle altitudes of 10 kft and 50 kft. Levels of combustible species within the plume shear layers are calculated in order to assess assumptions made in the base heating analysis.

Tucker, P. K.↗
NASA NTRS1993

A CFD validation roadmap for hypersonic flows

A roadmap for computational fluid dynamics (CFD) code validation is developed. The elements of the roadmap are consistent with air-breathing vehicle design requirements and related to the important flow path components: forebody, inlet, combustor, and nozzle. Building block and benchmark validation experiments are identified along with their test conditions and measurements. Based on an evaluation criteria, recommendations for an initial CFD validation data base are given and gaps identified where future experiments would provide the needed validation data.

Marvin, Joseph G.↗
NASA NTRS1992

Images constructed from computed flow fields

A method for constructing interferograms, schlieren, and shadowgraphs from ideal- and real-gas, two- and three-dimensional computed flow fields is described. The computational grids can be structured or unstructured, and multiple grids are an option. The constructed images are compared to experimental images for several types of flow, including a ramp, a blunt-body, a nozzle, and a reacting flow. The constructed images simulate the features observed in the experimental images. They are sensitive to errors in the flow-field solutions and can be used to identify solution errors. In addition, techniques for obtaining phase shifts from experimental finite-fringe interferograms and for removing experimentally induced phase-shift errors are discussed. Both the constructed images and calculated phase shifts can be used for validation of computational fluid dynamics (CFD) codes.

Yates, Leslie A.↗
NASA NTRS1993

Numerical study of shock-induced combustion in methane-air mixtures

The shock-induced combustion of methane-air mixtures in hypersonic flows is investigated using a new reaction mechanism consisting of 19 reacting species and 52 elementary reactions. This reduced model is derived from a full kinetic mechanism via the Detailed Reduction technique. Zero-dimensional computations of several shock-tube experiments are presented first. The reaction mechanism is then combined with a fully implicit Navier-Stokes computational fluid dynamics (CFD) code to conduct numerical simulations of two-dimensional and axisymmetric shock-induced combustion experiments of stoichiometric methane-air mixtures at a Mach number of M = 6.61. Applications to the ram accelerator concept are also presented.

Yungster, Shaye↗
NASA NTRS1993

Physical properties of the benchmark models program supercritical wing

The goal of the Benchmark Models Program is to provide data useful in the development and evaluation of aeroelastic computational fluid dynamics (CFD) codes. To that end, a series of three similar wing models are being flutter tested in the Langley Transonic Dynamics Tunnel. These models are designed to simultaneously acquire model response data and unsteady surface pressure data during wing flutter conditions. The supercritical wing is the second model of this series. It is a rigid semispan model with a rectangular planform and a NASA SC(2)-0414 supercritical airfoil shape. The supercritical wing model was flutter tested on a flexible mount, called the Pitch and Plunge Apparatus, that provides a well-defined, two-degree-of-freedom dynamic system. The supercritical wing model and associated flutter test apparatus is described and experimentally determined wind-off structural dynamic characteristics of the combined rigid model and flexible mount system are included.

Dansberry, Bryan E.↗
NASA NTRS1993

Computational analysis of Variable Thrust Engine (VTE) performance

The Variable Thrust Engine (VTE) of the Orbital Maneuvering Vehicle (OMV) uses a hypergolic propellant combination of Monomethyl Hydrazine (MMH) and Nitrogen Tetroxide (NTO) as fuel and oxidizer, respectively. The performance of the VTE depends on a number of complex interacting phenomena such as atomization, spray dynamics, vaporization, turbulent mixing, convective/radiative heat transfer, and hypergolic combustion. This study involved the development of a comprehensive numerical methodology to facilitate detailed analysis of the VTE. An existing Computational Fluid Dynamics (CFD) code was extensively modified to include the following models: a two-liquid, two-phase Eulerian-Lagrangian spray model; a chemical equilibrium model; and a discrete ordinate radiation heat transfer model. The modified code was used to conduct a series of simulations to assess the effects of various physical phenomena and boundary conditions on the VTE performance. The details of the models and the results of the simulations are presented.

Giridharan, M. G.↗
NASA NTRS1993

An experimental study of characteristic combustion-driven flows for CFD validation

The application of laser-based diagnostic techniques has become commonplace to a wide variety of combustion problems. New insights into combustion phenomena at a level previously unattainable has been made possible by non-intrusive measurements of velocity, temperature, and species. However, due to the adverse conditions which exist inside rocket engines, relatively few studies have addressed these combustion environments. The high pressure, high speed, combusting environment in a rocket engine prohibits the application of several measurement techniques. However, in the rocket community, there is a critical need for rocket flow field data to validate computational fluid dynamic (CFD) codes. Currently at Penn State, there is an effort to obtain flowfield measurements inside a rocket engine. Velocity measurements have been made inside the combustion chamber of a uni-element (shear coaxial injector) optically accessible rocket chamber at several axial locations downstream of the injector. These measurements, combined with future measurements, will provide benchmark data for CFD code validation.

Pal, S.↗
NASA NTRS1994

Radiative properties of molecular nitrogen ions produced by helium Penning ionization and argon effects

The development of hypersonic aerospace vehicles requires a better understanding on the thermal and chemical nonequilibrium kinetics of participating species in shock layers. The computational fluid dynamic (CFD) codes developed for such flowfields overestimate the radiation in the spectral region of 300 - 600 nm. A speculation for this overestimation is that inclusion of Ar, CO2, and H2O at the upper atmosphere flight region makes a significant impact on radiative kinetics of molecular nitrogen ions. To define the effects of minority species on the radiative kinetics of N2(+), an experimental setup was made by using the helium Penning ionization. The vibrational and rotational temperature were measured by mapping the vibrational and rotational distributions of N2(+) emission with high spectroscopic resolution and absolute intensity measurements. Measured vibrational temperatures were in the range from 18,000 to 36,000 K, and rotational temperatures were in the range from 300 to 370 K. The irradiance of 391.44 nm line and rotational and vibrational temperatures were analyzed to define argon and CO2 effects on the N2(+) emission. When Ar or CO2 is injected with N2, the rotational temperature did not change. The irradiances were reduced by 34 percent and 78 percent for the 50 percent of mixture of Ar and CO2, respectively. The vibrational temperatures were increased by 24.1 percent and 82.9 percent for the 50 percent of mixture of Ar and CO2, respectively. It appears that there are no significant effects from small concentrations of Ar and CO2 at the upper atmosphere flight region.

Miller, George, III↗