David W. Bogdanoff

Shock Radiation Measurements for Mars Aerocapture Radiative Heating Analysis

NASAs In-Space Propulsion Technology program is supporting the development of shock radiation transport models for aerocapture missions to Mars and Venus. Phenomenological models of nonequilibrium shock radiation will be incorporated into high-fidelity flowfield computations used to predict the aerothermal environments for a Mars or Venus aerocapture entry vehicle. These models are validated with shock radiance measurements obtained at flight-relevant conditions. A comprehensive test series in the NASA Ames Electric Arc Shock Tube facility at a representative freestream condition was recently completed. The facilitys optical instrumentation enabled spectral measurements of shocked gas radiation from the vacuum ultraviolet to the near infrared. The instrumentation captured the nonequilibrium postshock excitation and relaxation dynamics of dispersed spectral features. A description of the shock tube facility, optical instrumentation, and examples of the test data are presented.

Analysis and Model Validation of Shock Layer Radiation in Air

This paper analyzes the shock layer radiative heating environment for a large entry vehicle on a lunar return trajectory. Modeling results show that much of the shock layer plasma is in local thermodynamic equilibrium (LTE) and is not optically thin. The ionization level is generally high (15%) and the air is almost fully dissociated. A significant amount of vacuum ultraviolet (VUV) radiation is produced due to bound-bound and bound-free transitions of N and O atoms. The sensitivity of total radiation to Stark broadening, which dominates over other line broadening mechanisms, is quantified. The latter part of this paper reports the status of ongoing validation of the current radiation models with measurements in the Electric-Arc Shock Tube (EAST) facility at NASA Ames Research Center. Model predictions are compared with the calibrated radiation spectra measured in the equilibrium portion of the shock layer at 0.3 Torr. The reasons for discrepancy between model and measurements are also discussed with possible hypotheses presented for further investigation.

Shock-Tube Measurement of Nitridation Coefficient of Solid Carbon

The coefficient of reaction of nitrogen atoms with solid carbon to form gaseous cyanogen CN is measured in a shock tube. A stream of highly dissociated nitrogen is produced in a shock tube and passed over a grid of metal wire coated with carbon. The radiation emitted by the wake of the wire grid is observed with a monochromator set at 386 nm, where CN is known to radiate strongly. From the intensity of the CN radiation, the concentration of CN in the wake is determined. The concentration of nitrogen atoms in the stream are calculated by integrating conservation equations. From these concentration values, the fraction of the collisions of nitrogen atoms producing CN, that is, the reaction coefficient, is deduced. The results show that the reaction coefficient is about 0.3 at both 300 and 1100 K. The uncertainties in the results are described, and improvements are proposed.

Modeling and Experimental Validation of CN Radiation Behind a Strong Shock Wave

Assessment of nonequilibrium thermochemical models for shock layer radiation in N2/CH4 mixtures is presented via comparisons with spectrally and temporally resolved intensity measurements from a set of shock tube experiments. The experiments were carried out at the Electric Arc Shock Tube (EAST) facility at NASA Ames Research Center in a rarified environment [13.3-133.3 Pa (0.1 and 1 Torr)] representative of the peak heating conditions of a Titan aerocapture trajectory (5-9 km/s). The baseline model that assumes a Boltzmann population of the CN excited states consistently overpredicts the shock layer radiation intensity. A non-local collisional radiative model that solves a simplified master equation and includes radiative transport and non-local absorption in the shock tube is presented. The proposed model improves the prediction of the nonequilibrium radiation overshoot peak, but still underpredicts the intensity decay rate in the low pressure case. Further analysis suggests possible reasons for the remaining disagreement, the most likely being a slow CN consumption in the current chemical kinetics model in the intensity fall-off region.

Comparisons of Air Radiation Model with Shock Tube Measurements

This paper presents an assessment of the predictive capability of shock layer radiation model appropriate for NASA s Orion Crew Exploration Vehicle lunar return entry. A detailed set of spectrally resolved radiation intensity comparisons are made with recently conducted tests in the Electric Arc Shock Tube (EAST) facility at NASA Ames Research Center. The spectral range spanned from vacuum ultraviolet wavelength of 115 nm to infrared wavelength of 1400 nm. The analysis is done for 9.5-10.5 km/s shock passing through room temperature synthetic air at 0.2, 0.3 and 0.7 Torr. The comparisons between model and measurements show discrepancies in the level of background continuum radiation and intensities of atomic lines. Impurities in the EAST facility in the form of carbon bearing species are also modeled to estimate the level of contaminants and their impact on the comparisons. The discrepancies, although large is some cases, exhibit order and consistency. A set of tests and analyses improvements are proposed as forward work plan in order to confirm or reject various proposed reasons for the observed discrepancies.

Theoretical Performance of a Magnetohydrodynamic-Bypass Scramjet Engine with Nonequilibrium Ionization

The theoretical performance of a scramjet propulsion system in which a magnetohydrodynamic (MHD) energy bypass scheme is operated in a nonequilibrium ionization environment is evaluated. In the MHD generator, the incomingairflow at a temperature too low to produce the required ionization is seeded with cesium and ionized by the use of an unspecified external power source. The resulting nonequilibrium environment is described using a two-temperature model. The accelerator is operated with equilibrium ionization. The expansion through the nozzle is calculated with account taken of finite rate reactions, and the boundary layer is assumed to be fully turbulent. The results show that the required external power is of the same order of magnitude as that of the power generated in the MHD generator. An MHD energy bypass propulsion scheme looks to be more promising with the equilibrium ionization method than with the nonequilibrium ionization method.

Theoretical Performance of Frictionless Magnetohydrodynamic-Bypass Scramjets

Theoretical performance calculation is made of a scramjet propulsion system incorporating a magnetohydrodynamic (MHD)-energy-bypass scheme. The MHD generator upstream of the combustion chamber slows down the e ow so that the Mach number at the entrance of the combustion chamber is kept below a specie ed value. The MHD accelerator downstream of the combustion chamber accelerates the e ow, expending the electrical power produced by the generator. The e ow is seeded with potassium or cesium, and theMHD devices operate as Faraday machines. Friction is neglected, and chemical equilibrium isassumed everywhere except in the nozzle downstream of the freezing point. The calculation shows that the MHD-bypass scheme can improve specie c impulse over that of a conventional scramjet at e ight speeds over 3.5 km/s. At speeds below about 6 km/s, the calculated specie c impulse can be higher than that of a typical rocket engine. Consequently, the MHD-bypass scheme can extend the operational range or improve the performance of a conventional scramjet engine.

Improving the Performance of Two-Stage Gas Guns By Adding a Diaphragm in the Pump Tube

Abstract Herein, we study the technique of improving the gun performance by installing a diaphragm in the pump tube of the gun. A CFD study is carried out for the 0.28 in. gun in the Hypervelocity Free Flight Radiation (HFF RAD) range at the NASA Ames Research Center. The normal, full-length pump tube is studied as well as two pump tubes of reduced length (∼75% and ∼33% of the normal length). Significant improvements in performance are calculated to be gained for the reduced length pump tubes upon the addition of the diaphragm. These improvements are identified as reductions in maximum pressures in the pump tube and at the projectile base of ∼20%, while maintaining the projectile muzzle velocity or as increases in muzzle velocity of ∼0.5 km/sec while not increasing the maximum pressures in the gun. Also, it is found that both guns with reduced pump tube length (with diaphragms) could maintain the performance of gun with the full length pump tube without diaphragms, whereas the guns with reduced pump tube lengths without diaphragms could not. A five-shot experimental investigation of the pump tube diaphragm technique is carried out for the gun with a pump tube length of 75% normal. The CFD predictions of increased muzzle velocity are borne out by the experimental data. Modest, but useful muzzle velocity increases (2.5 – 6%) are obtained upon the installation of a diaphragm, compared to a benchmark shot without a diaphragm.

Reactivation and upgrade of the NASA Ames 16-Inch Shock Tunnel - Status report

The NASA Ames 16-Inch Shock Tunnel has been reactivated after seventeen years of inactivity. In the years before deactivating the facility, it was operated at enthalpies of 4,700 J/gm and pressures up to 260 atm or at enthalpies of 1900 J/gm over a wide pressure range. Since reactivating, the facility has been operated at enthalpies up to 12,000 J/gm and pressures up to 408 atm. The present paper describes the steps taken in upgrading the facility and summarizes the currently achievable conditions. The selection of the driver gas, the steps taken to improve the driver burn, and the diaphragm opening techniques are described. The pressure and heat flux instrumentation, the optical diagnostics and the data acquisition system are also described.

Free-Flight Measurements of Convective Heat Transfer in Hypersonic Ballistic-Range Environments

The NASA-Ames Hypervelocity Free-Flight Aerodynamic Facility is a ground-based facility for real-gas aerothermodynamic testing, offering the unique ability to independently vary the velocity, the effective altitude (static pressure), and the test gas composition. A technique is demonstrated to determine quantitative, global convective heat transfer rates from high-speed thermal images of hypersonic projectiles in flight in this facility. Measurements were made in air and in nitrogen on titanium alloy hemispheres at velocities up to 4.5 km/sec. Results compared within ±10% of published stagnation-point heat transfer rate measurements and with established engineering correlations. Real-gas NavierStokes computations are in agreement with the measurements.