Shock Layer Flow Research Articles

Numerical Analysis of Aerodynamic Thermal Properties of Hypersonic Blunt-Nosed Body with Angles of Fire

A hypersonic electromagnetic railgun projectile undergoes severe aero-heating with an increase in altitude. The purpose of this study was to investigate the characteristics of the shock layer flow field as well as the thermal environment of the blunt body wall of a hypersonic electromagnetic railgun projectile at different launching angles. The two-temperature model considers the thermal nonequilibrium effect and is introduced into the Navier–Stokes (N-S) equation, and it is solved using the finite volume method (FVM). The reliability of the calculation model in terms of thermal properties and composition production was verified against a blunted-cone-cylinder–flare (HB-2) test case. The surface temperature of the hypersonic blunt projectile was simulated using a radiation balance wall boundary. The thermal characteristics at the emission angles α = 60° and α = 45° were checked within an altitude range of 0–70 km, including the nonequilibrium effect, reaction heat release, aerodynamic heat flux, and wall temperature. The results show that the translational rotational temperature is higher than the vibrational electronic temperature, and the thermal nonequilibrium effect increases with an increase in altitude. Comparing the two launching angles, the nonequilibrium degree and reaction heat release at α = 60° were higher than those at α = 45°. The rates of exothermic reaction decreased with an increase in altitude. The heat flux along the wall of the generatrix decreased sharply from the stagnation point. With an increase in altitude, the heat flux dropped sharply from 7 MW/m2 at H = 0 km to approximately 2 MW/m2 at H = 70 km. The wall temperature distribution was similar to the heat flux distribution; however, the surface temperature decreased less rapidly than the heat flux.

Mathematical Simulation of Nonequilibrium Shock Layer Flow around a Rotating Body

Mathematical Simulation of Nonequilibrium Shock Layer Flow around a Rotating Body

Assessment of nonequilibrium air-chemistry models on species formation in hypersonic shock layer

The present study aims to assess the performance of chemical rate models for species formation in the reacting shock layer flows over the blunt-cone hypersonic vehicles. Three 11-species nonequilibrium chemical models for air, Gupta 90, Park 93 and Ozawa’s modified models, are assessed. Two controlling temperature expressions (T0.5Tv0.5 and T0.7Tv0.3) for dissociation reactions are taken into account in these chemical kinetic models. To further examine the performance of these models for predictions of nonequilibrium effects and species formations in the shock layer, two typical flight cases are adopted: (1) the NO formation in reacting flows over the Bow-shock Ultraviolent (BSUV) vehicle at Mach number 17.7, and (2) the electron formation over the Radio Attenuation Measurements (RAM) C-II vehicle at Mach number 23.9 and 25.9. Firstly, comparisons of interested parameters between the computed results in different rate models and available reference data are carried out. Secondly, the reasons for the difference of species formations in these models are discussed. Results show that both the chemical rate model and the weight factor of the controlling temperature have a distinctive influence on species concentration and distribution in the shock layer. The weight factor determines the level of the vibrational-electronic temperature and the reaction release heat in nonequilibrium processes. With the increasing of the weight factor, the NO concentration increases and the electron density decreases in the same rate model. The spectral integration within the wavelengths of 205–255 nm shows that the prediction accuracy of the Park-0.5 and Ozawa-0.7 models is relatively high. Numerical results also indicate that the Ozawa-0.7 model may be an all-round model to predict electron formation in the shock layer.

Experimental evidence of the impact of radiation coupling on binary scaling applied to shock layer flows

Binary scaling is a similitude law used to study the aerothermodynamics of hypersonic vehicles in ground-based high-enthalpy facilities. It enables the duplication of shock layer flows in the vicinity of the stagnation point, including binary chemistry and nonequilibrium processes, over length scales that are practical for experimental testing. It is built on the assumption of a flow devoid of radiation coupling, which drastically narrows down the envelope of flows for which binary scaling can be used. Indeed, if it is strong enough, radiative heat transfer will cause a substantial amount of energy to leak out of the shock layer to the free-stream and other flow regions. As demonstrated in this paper, the strength of that coupling will increase as the length-scale of the flow increases, impacting other flow properties such as its chemical composition or temperature. The resulting impact on macroscopic features of the flow are for example a reduction of the wall heat flux (both convective and radiative) or of the shock standoff distance. These side impacts are identified with experimental measurements on the shock standoff distance in an expansion tube with CO2–N2 mixture flows representative of the venusian atmosphere over cylinders.

Distinctive features of the shock layer flow structure in conical gas flows

The results of a numerical and experimental investigation of the flow structure in symmetric and nonsymmetric flows around V-wings with attached shocks on the leading edges are presented. Emphasis is placed on the appearance of new critical points, including vortex Ferri singularities, in the shock layer and transformations in the flow structure with increase in the angles of attack and yaw. In particular, it is established that the flow structure in the plane of symmetry of the flow around V-wings without yaw, which involves the Mach-type shock configuration, undergoes a jumpwise variation with increase in the angle of attack. Additionally to one Ferri singularity of the node type located at the corner point of the transverse wing contour in the plane of symmetry of the flow, there arise two more critical points, those of flow divergence and convergence. The latter point is the second Ferri singularity; it is located nearer to the bridge-shaped shock of the Mach-type shock system and can be of both the node and the saddle type. In the latter case there appear two vortex Ferri singularities located at the vertices of the contact discontinuities proceeding from this critical point on both sides of the plane of symmetry. Certain data on the position of the critical points relative to the wing contour bend are presented as functions of the wing geometry, together with the transformation of the topological shock-layer flow pattern in the presence of yaw. The comparison of the results calculated within the framework of the Euler model with the experimental data on the shock-layer flow structure obtained using a special optical method for visualizing conical flows showed their good agreement.

Modeling of Unsteady Shock Tube Flows Using Direct Simulation Monte Carlo

The simulation of the UV spectra of moderate Mach number flows is the strongest test possible of the complex thermochemical models in hypersonic flows. To take advantage of low-density ground-based measurements, however, one needs to be able to model unsteady flows using kinetic particle approaches. The direct simulation Monte Carlo approach is used in this paper to model the thermal relaxation in nonreacting nitrogen shocks corresponding to the data and the higher Mach number air shocks and radiation of previous research. In the first case, it is found that better agreement with experimental data can be achieved when assuming a constant rotational collision number instead of a temperature-dependent one. For the second data set, the nonequilibrium NO UV radiation from unsteady reflected air shocks are modeled and compared with experimental data. The Larsen–Borgnakke and total collision energy models are used to simulate the internal energy relaxation and chemistry of five chemical species (, , N, O, and NO), respectively. Unlike steady shock layer flows, the long period of equilibration behind the shock wave occurs and potential factors that prevent complete thermochemical equilibrium are explored. The main goal of this paper, however, is to compare the spectrally averaged peak shock radiation with data. Because there are insufficient collisions to assume a Boltzmann distribution of the NO electronic states, a set of excitation/deexcitation mechanisms are proposed and the quasi-steady-state assumption is used in calculating the neutral-impact excitation of NO and the associated radiation. The calculated spectrally averaged peak radiations in the shock are in good agreement with the experimental data of Gorelov for shock speeds below , but are lower at higher shock speeds.

Computational Analysis of Radiative Heat Loading on Hypervelocity Reentry Vehicles

Space exploration missions may include atmospheric entries at very high velocities (for example, sample return or entries from hyperbolic orbits) or into an atmosphere with highly radiative constituents (for example, Mars or Titan). The effects of gas radiation in the shock layer and wake flows may become significant at such flight conditions; hence, an accurate computational fluid dynamics analysis of these flows requires adequate modeling of the local absorption and emission properties and the energy transport by radiation. The wide range of numerical techniques for the integration of radiation in entry vehicle simulations includes Monte Carlo models, which enable accurate and computationally efficient predictions of radiative energy transport in participating (absorbing and emitting) media for complex geometries. Reentry type flow properties are characterized by a large variation of local absorption and emission properties in the spectral range. Therefore, and due to the absence of general averaging rules, a fine spectral discretization is needed for practical computations. Here, binning methods offer the potential to reduce the required resolution while maintaining a high level of accuracy. The application of a Monte Carlo radiation transport code using high spectral resolution and a loose coupling approach with the applied flow solver was found to give good agreement with the available heating rate measurements of the Second Flight Investigation of Reentry Environment flight experiment.

Effects of thermal boundary conditions on the hypersonic shock layer flow over a blunt body

This paper presents a numerical study of the effects of thermal boundary conditions on the shock layer flow over a spacecraft entering a planetary atmosphere at hyper-velocity, and on the resulting convective heat fluxes to its heat shield surface. This study is motivated by the fact that spacecraft heat shields are covered with materials that ablate to reduce the spacecraft wall temperature. The ablation is not taken into account, but the effects of changes in the temperature of the vehicle surface are thoroughly studied. The results show that the flow field solutions within the shock layer and the convective heat fluxes are sensitive to the wall thermal boundary conditions. A computation of the flow field solution for an adiabatic wall is also performed. In this case, the gas recovery temperature along the wall increases as the gas moves away from the stagnation point. Computation convergence is also analysed.

Observation of an Ablating Surface in Expansion Tunnel Flow

The observation of an ablating surface in expansion tunnel flow was reported. A one-dimensional (1-D) semi-infinite analysis was performed using an empirical estimation of the stagnation-point heat flux. A surface temperature change for epoxy of 178 K in 50 μs was calculated, which is sufficient for the epoxy coating to commence ablation during the steady test period. The 10% temperature penetration depth is 6 μm in 50 μs. The luminosity from substantial portions of the model and shock-layer flow was visualized using a Shimadzu HPV1 high-speed charge-coupled device (CCD) video camera recording at 500 kfps with a 1 μs exposure time. The uniform image response over the axisymmetric model implies the shock-layer gas irradiance is high immediately behind the shock and decreases rapidly as the model surface is approached. The measurements showed that use of an epoxy coating results in greatly increased CN and C2 radiation, much greater than when no coating is employed.

Tomographic reconstruction of shock layer flows

The tomographic reconstruction of supersonic flows faces two challenges. Firstly, techniques used in the past, such as the direct Fourier method (DFM) (Gottlieb and Gustafsson in On the direct Fourier method for computer tomography, 1998; Morton in Tomographic imaging of supersonic flows, 1995) or various backprojection (Kak and Slaney in Principles of computerized tomographic imaging, vol. 33 in Classics in Applied Mathematics, 2001) techniques, have only been able to reconstruct areas of the flow which are upstream of any opaque objects, such as a model. Secondly, shock waves create sharp discontinuities in flow properties, which can be difficult to reconstruct both in position and in magnitude with limited data. This paper will present a reconstruction method, matrix inversion using ray-tracing and least squares conjugate gradient (MI-RLS), which uses geometric ray-tracing and a sparse matrix iterative solver (Paige and Saunders in ACM Trans. Math. Softw. 8(1):43–71, 1982) to overcome both of these challenges. It will be shown, through testing with a phantom object described in tomographic literature, that the results compare favourably to those produced by the DFM technique. Finally, the method will be used to reconstruct three-dimensional density fields from interferometric shock layer images, with good resolution (Faletič in Tomographic reconstruction of shock layer flows, 2005).

G407 炭素系アブレーションを伴う融合衝撃層の多温度解析(2)(GS-4-2 飛翔体,円柱,一般セッション)

Numerical simulation of nonequilibrium shock-layer flows was performed for the reentry conditions of MUSES-C with carbonaceous ablation effects by an extended merged shock-layer theory based on a 4-temperature and 16 chemical species model. The numerical calculations assumed a continuum approach and the system of governing equations constituted a two-point boundary value problem on a stagnation-point streamline, which was solved numerically by the SAR method. The effects of ablation on temperatures and flowfields were discussed in detail. It was found that the ablation from a body surface had significantly influence on various physico-chemical values in the merged shock-layer.

ダストを含む超音速流中の鈍頭物体衝撃層に関する数値解析

In this paper, supersonic dusty shock-layer flow over a blunt body is numerically studied by the Eulerian and Lagrangian hybrid approach. For the gas phase, the axisymmetric Navier-Stokes equations are solved by the finite difference method, while the particle phase is analyzed by the method of particle tracing. The bounce of particles at the body surface is considered. The gas-particle interaction is taken into account by evaluating the momentum and energy exchange between the particle and the ambient fluid. The numerical results show that the presence of particles in the flow causes the reduction in the shock-layer thickness and the augmentation of the wall pressure and wall heating rate. These effects strongly depend on the size of the particles.

Coherent Anti-Stokes Raman Scattering Measurements and Computational Modeling of Nonequilibrium Flow

An experimental study is conducted in a high-enthalpy continuous facility to investigate the nonequilibrium shock-layer airflow induced by a two-dimensional body. The rotational and vibrational temperatures of nitrogen as well as the number density are measured by recording the rovibrational spectra of molecular nitrogen using coherent anti-Stokes Raman scattering spectroscopy. Measurements are performed in the freestream and within the shock layer induced by a disk model. Difficulties specific to the application of an optical technique to a high enthalpy flow are discussed, and flows parameters are given. The nozzle and shock-layer flows are modeled numerically using a nonequilibrium Navier-Stokes solver. Good agreement is achieved between the experimental and calculated results.

Three-dimensional supersonic turbulent flows past conical bodies

The results of the numerical simulation of supersonic three-dimensional flow past sharp-nosed cones with circular and elliptic cross-sections in the turbulent shock-layer flow regime are presented. The calculations are performed in the local conical approximation using the system of Reynolds equations and the differential one-equation turbulence model. The numerical solutions are obtained by means of an implicit constant-direction finite-difference scheme. The emphasis is placed on the investigation of the transverse flow separation and the flow features associated with the turbulent flow regime.

Interaction of Spalled Particles with Shock Layer Flow

The influence of the solid particles injected into the shock layer by the spaliation of a carbonaceous ablating heatshield on the shock layer flow is investigated theoretically. The equation of motion of a solid particle is integrated using a polynomial expansion. Assuming a Gaussian distribution for the initial masses and velocities of the particles, the rates of production of CO, CN, C 3 , and turbulence energy are calculated by integrating those production rates for one particle over the mass and velocity distributions. The results for the environment of Stardust Earth reentry show that these production rates decay approximately exponentially in the normal direction from the ablating wall. Correlation formulas for the magnitudes and slopes of this decay pattern are derived. The lower limits of particle size and initial velocity are estimated for a carbon-phenolic material from a spectroscopic result obtained in an arcjet wind-tunnel test

Numerical Simulation of Laser-Induced Fluorescence Imaging in Shock-Layer Flows

Planar laser-induced fluorescence (PLIF) images of nitric oxide in hypersonic flow over a wedge and a hemisphere are compared with a theoretical PLIF model. The theoretical PLIF images are based on computational fluid dynamics (CFD) models including a perfect-gas model and a nonequilibrium chemistry model. Two-dimensional maps of the flow parameters generated by the CFD are used to predict the theoretical PLIF images, including the effects of collisional quenching. We find good agreement between the model and the experimental measurements. We explain how this method of computational flow imaging can be useful for designing experiments.

Numerical simulation of laser-induced fluorescence imaging in shock-layer flows

Numerical simulation of laser-induced fluorescence imaging in shock-layer flows

Spectroscopic Measurements of Shock-Layer Flows in an Arcjet Facility

Under two different test conditions, radiation emanating from the freestream and shock-layer flow over a 15.24-cm-diam blunt-body test article as measured in NASA Ames Research Center's 20 MW Aerodynamic Heating Facility Arcjet. The test gas was a mixture of air and argon. Spatially resolved emission spectra were obtained over a 2000-8900 A wavelength range using a charge-coupled device camera (1024 × 256 array) attached to a spectrograph. The optical system was calibrated using tungsten and deuterium radiation sources. Previously developed analytical tools were used to determine the following line-of-sight-averaged thermodynamic properties from the calibrated spectra: 1) rotational temperature of the freestream and 2) rotational and vibrational temperatures within the shock layer. Based on the variation in intensity of emission spectra along the stagnation streamline, shock standoff distance was determined. Two sets of data for each test condition were compared to evaluate the repeatability of measurements. Considering likely sources of errors, an uncertainty analysis was performed to estimate the error bounds of the determined properties

Computational Simulation of Emission Spectra from Shock-Layer Flows in an Arcjet Facility

This paper reports computational comparisons with experimental studies of a nonequilibrium bluntbody shock-layer e ow in a high-enthalpy arcjet wind tunnel at NASA Ames Research Center. The experimental data include spatially resolved emission spectra of radiation emanating from a shock layer formed in front of a e at-faced cylinder model. Multitemperature nonequilibrium codes are used to compute the conical nozzle e ow, supersonic jet, and shock-layer e ow. A line-of-sight (LOS) radiation code is employed to predict the spectra from the computed e owe eld. Computed LOS emission intensities are directly compared with the experimental data at several axial locations along the stagnation streamline. Various LOS-averaged e ow properties, such as vibrational and rotational temperatures, and species number densities, deduced from the experimental spectra, are compared with computed results. Comparisons provide an assessment of thermochemical equilibration processes in an arcjet shock layer.

Rotational and vibrational temperature measurements using CARS in a hypervelocity shock layer flow and comparisons with CFD calculations

Broadband single pulse coherent anti-Stokes Raman scattering (CARS) experiments employing a folded-box phase-matching geometry in a pulsed hypervelocity blunt body flow are presented. Rovibrational spectra of molecular nitrogen, produced in the freestream and within the shock layer at moderately high enthalpy (8.4 MJ/kg), are examined. Difficulties peculiar to the application of a single pulse optical technique to a high enthalpy pulsed flow facility are discussed and measurements of flow temperatures are presented. Theoretically calculated values for temperatures based upon algorithms used to determine freestream and shock layer conditions agree well with experimental measurements using the CARS technique. The measurements indicate that thermal non-equilibrium conditions exist within the freestream, and that near thermal equilibrium exists at the point of measurement within the shock layer. The comparison between the experiment and theory in the shock layer is improved by using the measured freestream temperatures as input to the shock layer computations.