Reentry Flows Research Articles

Study on the real gas effect on the windward side of an inflatable reentry capsule

In this paper, the reentry flow field of an inflatable reentry capsule was numerically simulated using the air 9-component chemical nonequilibrium model and the ideal gas hypothesis model. The difference in the calculation results of the two methods was investigated, the manifestation of the real gas effect was studied, and the reasons for the difference between the real gas effect of the inflatable reentry capsule and the real gas effect of the rigid reentry capsule were explored. The results show that compared with the ideal gas hypothesis, the real gas effect makes the shock wave position closer to the wall, the air temperature after the shock wave decreases, and the wall heat flux density decreases; at 83 km, the specific heat ratio of the gas after the shock wave is higher than 1.4, and the air undergoes a dissociation reaction, while at 73 km, the real gas effect is weaker, the specific heat ratio of the air after the shock wave remains at 1.4, and the air still exists in the form of molecules; the main reason for the difference in the strength of the real gas between the inflatable reentry capsule and the rigid reentry capsule in the same altitude range is that the drag-to-weight ratio of the inflatable reentry capsule is larger than that of the rigid reentry capsule, and the speed decreases faster after entering the atmosphere, and the speed is lower at the same altitude.

Re-entry rocket basic flow characteristics and thermal environment of different retro-propulsion modes

During the supersonic re-entry of multi-nozzle heavy rockets into the atmosphere, the basic flow state becomes increasingly complex due to the coupling effect between the retro-propulsion plumes and the freestream. A numerical method using the hybrid Reynolds-Averaged Navier-Stokes and Large Eddy Simulation (RES) method and discrete coordinate method is developed to accurately estimate the thermal environment. In addition, finite rate chemical kinetics is used to calculate the afterburning reactions. The numerical results agree well with wind tunnel data, which confirms the validity and accuracy of the numerical method. Computations are conducted for the heavy carrier rocket re-entry from 53.1 km to 39.5 km altitude with 180° angle of attack by using three different Supersonic Retro-Propulsion (SRP) modes. The numerical results reveal that these three SRP flow fields are all Short Penetration Models (SPM). As the re-entry altitudes decrease, both the plume-plume interaction and the plume-freestream interaction become weaker. The highest temperatures in the plume shear layers of the three SRP modes increase by 8.36%, 7.33% and 6.92% respectively after considering afterburning reactions, and all occur at a re-entry altitude of 39.5 km. As the rocket re-enters the atmosphere, the maximum heat flux on the rocket base plate of three SRP modes stabilizes at 290, 170 and 200 kW/m2 respectively, but the maximum heat flux on the side wall increases significantly. When the altitude declines to 39.5 km, the extreme heat flux of the three modes increase by 84.16%, 49.45% and 62.97% respectively compared to that at 53.1 km.

Numerical study on the non-equilibrium characteristics of high-speed atmospheric re-entry flow and radiation of aircraft based on fully coupled model

The strong coupling interactions of non-equilibrium flow, microscopic particle collisions and radiative transitions within the shock layer of hypersonic atmospheric re-entry vehicles makes accurate prediction of the aerothermodynamics challenging. Therefore, in this study a self-consistent non-equilibrium flow, collisional–radiative reactions and radiative transfer fully coupled model are established to study the non-equilibrium characteristics of the flow field and radiation of vehicle atmospheric re-entry. The comparison of the present calculation results with flight data of FIRE II and previous results in the literature shows a reasonable agreement. The thermal, chemical and excited energy level non-equilibrium phenomena are obtained and analysed for the different FIRE II trajectory points, which form the critical basis for studying the heat transfer and radiation. The non-equilibrium distribution of excited energy levels significantly exists in the post-shock and near-wall regions due to the rapid vibrational dissociation and electronic under-excitation, as well as the wall catalytic reactions. The analysis of stagnation-point heating of FIRE II illustrates that the translational–rotational convection and the dissociation component diffusion play key roles in the aerodynamic heating of the wall region. The spectrally resolved radiative intensity in the entire flow field indicates that the vacuum ultraviolet radiation caused by the high-energy nitrogen atomic spectral lines makes the main contribution to the radiative transfer. Finally, it is found that the non-equilibrium flow–radiation coupling effect can exacerbate the excited energy level non-equilibrium, and further affect the gas radiative properties and radiative transfer. This fully coupled study provides an effective method for reasonable prediction of atmospheric re-entry flow and radiation fields.

Data-Driven Modeling of Hypersonic Reentry Flow with Heat and Mass Transfer

The entry phase constitutes a design driver for aerospace systems that include such a critical step. This phase is characterized by hypersonic flows encompassing multiscale phenomena that require advanced modeling capabilities. However, because high-fidelity simulations are often computationally prohibitive, simplified models are needed in multidisciplinary analyses requiring fast predictions. This work proposes data-driven surrogate models to predict the flow and mixture properties along the stagnation streamline of hypersonic flows past spherical objects. Surrogate models are designed to predict the velocity, pressure, temperature, density, and air composition as functions of the object’s radius, velocity, reentry altitude, and surface temperature. These models are trained with data produced by numerical simulation of the quasi-one-dimensional Navier–Stokes formulation and a selected Earth atmospheric model. Physics-constrained parametric functions are constructed for each flow variable of interest, and artificial neural networks are used to map the model parameters to the model’s inputs. Surrogate models were also developed to predict surface quantities of interest for the case of nonreacting or ablative carbon-based surfaces, providing alternatives to semiempirical correlations. A validation study is presented for all the developed models, and their predictive capabilities are showcased along selected reentry trajectories of space debris from low Earth orbits.

The Meshless Direct Simulation Monte Carlo method

A new numerical simulation method called the Meshless Direct Simulation Monte Carlo (DSMC) Method is presented in this article for solving a rarefied flow field problem. It uses a discrete points group instead of using a conventional mesh to discretize the computational domain. The meshless theory is used to solve the governing equations of the rarefied flow field. Numerical challenges related to mesh distortion and low accuracy encountered in conventional mesh technology are eliminated. The meshless technology used in the DSMC method constructs the molecular cloud structure and the virtual volume to establish the methodology of molecular fast search, molecular collision pair number calculation, and molecular cloud sampling. The numerical results of a two-dimensional flow around a cylinder, a nozzle flow, a three-dimensional flow around a sphere, and a Mars probe re-entry flow are presented to demonstrate the feasibility, accuracy, and robustness of the Meshless DSMC method. This study provides the basis for meshless technology development in the field of rarefied flow.

Uniform impingement heat transfer distribution in corrugated channel with an anti-crossflow-wing

For a further improvement of impingement jet array heat transfer in a corrugated channel, a wing-shaped structure design is proposed to weaken crossflow degradation. The naphthalene sublimation method was used to measure local heat/mass transfer distributions in the present study. Pressure drop characteristics were examined through pressure measurements. Detailed flow structures were analyzed through numerical simulations. In order to validate the superiority of the current wing structure, it was compared with the impingement heat transfer values from a typical rectangular channel and a conventional corrugated channel. The hole diameter-based Reynolds number was fixed to 10,000. The impingement distance and spacing between holes were h/d = 1 and s/d = 5, respectively. The hole patterns were designed to be in-lined in all cases. As the wing structure was added to the corrugated channel, the impingement supply flows of downstream holes became axisymmetric. The wing structure noticeably suppressed the re-entry flow from the corrugated channel back into the main/impingement channel, particularly under the downstream row of impingement holes. As a result, the overall area-averaged Sherwood number increased by about 6.4%, and the heat transfer non-uniformity in the streamwise direction was significantly reduced.

New insights on the ablation mechanism of silicon carbide in dissociated air plasmas

Thermal protection system (TPS) applied in space vehicle is mainly composed of silicon carbide (SiC) and related components. When exposed in re-entry flows, the SiC based TPS may suffer from severe ablation if the surface temperature is beyond a critical threshold (generally ≥1600°C). Unveiling the ablation mechanism of SiC is thus vitally important to define the borders of service and to help optimize the TPS structures in re-entry environments. Motivated by this, this work uses an inductively coupled plasma wind tunnel to generate the dissociated air plasmas and explores the ablation behaviors of several SiC materials (including SiC ceramic and SiC fiber reinforced SiC matrix composites). Some new insights are acquired based on the detailed analysis of the surface characteristics prior to and after ablations. First, the role of surface roughness on the critical ablation temperature of the SiC based composite is put forward. Second, a ‘from-point-to-surface’ ablation process due to the surface roughness during the short ‘temperature jump’ is proposed, which is an important supplementary to the existing ablation model of SiC; Finally, a smoother surface with less defects is concluded to efficiently enhance the ablation resistance of the SiC based TPS in re-entry environments.

Abstract No. 500 Evolution of the endovascular management of chronic type B aortic dissection: 20-year retrospective review

Chronic type B aortic dissection (cTBAD) is frequently complicated by aneurysmal degeneration and has been treated with TEVAR in high risk-patients. However, there are reports of high reintervention rates secondary to persistent false lumen perfusion from Type R re-entry flow.1 Two strategies to address this issue include false lumen embolization, commonly by custom made candy-plug devices (CP), and disruption of the thickened intimal flap to create a common aortic channel for stent-graft landing, commonly by laser-assisted septotomy (LAS).

Analysis of the Ambipolar Diffusion Approximation in Hypersonic Ionizing Reentry Flows

Simulations of the flow surrounding the Stardust Return Capsule at Mach 40 at various altitudes were used to test the validity of the ambipolar diffusion approximation. Results of direct simulation Monte Carlo (DSMC) simulations at 71 km altitude show that atomic N and O are dominant in the shock layer. The highest concentration of ions produced in the shock corresponds to , followed by , , , and . The only charged species present in the wake are , , and electrons, as , , and are formed and then depleted in the bow shock. At 100 km altitude, the influence of charged species is negligible. Simulations were also conducted in which the ambipolar diffusion approximation was not applied, and, instead, electrons were transported at the DSMC velocity. There was a net negative charge upstream from the bow shock and net positive charge in the plasma sheath adjacent to the vehicle. The magnitude of the electric field needed to maintain quasi neutrality is discussed. Results indicate that, for practical reentry flow conditions, the Debye length in the bow shock and wake is orders of magnitude smaller than the mean free path, and that invoking the ambipolar diffusion approximation is valid.

Effects of air properties on flow filed characteristics of inflatable reentry decelerator

Air properties are very important for predicting the flow filed characteristics of aircraft. At present, Sutherland formula is always used to predict the flow filed characteristics of inflatable reentry decelerator. However, high temperature can be observed in high speed reentry flow around inflatable reentry decelerator, thus the Sutherland formula is not suitable for reentry flow simulation of inflatable reentry decelerator. In order to investigate the effect of the air properties on the flow filed characteristics of inflatable reentry decelerator, and provides the theoretical references for the inflatable reentry decelerator design, the Gupta fitting formula and Sutherland formula are used to describe the air properties, respectively, for studying the flow filed characteristics of inflatable reentry decelerator. The results show that the reentry flow temperature with Gupta fitting formula used is lower than that with Sutherland formula used, and the temperature difference is more obvious in the high inflow Mach number and temperature cases. The results in the windward of inflatable reentry decelerator show that the Sutherland formula underestimates the heat flux and pressure, and the heat flux difference is much more obvious. Such as case 5, the peak heat flux obtained by Gupta fitting formula is by 29% higher than that obtained by Sutherland formula.

Three-Temperature Model Applied to Thermochemical Non-Equilibrium Reentry Flows in 2D—Seven Species

In this work, a study involving the fully coupled Euler and Navier-Stokes reactive equations is performed. These equations, in conservative and finite volume contexts, employing structured spatial discretization, on a condition of thermochemical non-equilibrium, are analyzed. High-order studies are accomplished using the Spectral method of Streett, Zang, and Hussaini. The high enthalpy hypersonic flows around a circumference, around a reentry capsule, along a blunt body, and along a double ellipse in two-dimensions are simulated. The Van Leer, Liou and Steffen Jr., and Steger and Warming flux vector splitting algorithms are applied to execute the numerical experiments. Three temperatures, which are the translational-rotational temperature, the vibrational temperature, and the electron temperature, are used to accomplish the numerical comparisons. Excellent results were obtained with minimum errors inferior to 6.0%. The key contribution of this work is the correct implementation of a three temperature model coupled with the implementation of three algorithms to perform the numerical simulations, as well the description of energy exchange mechanisms to perform more realistic simulations.

Analysis of mass diffusion theory and models for high-temperature multi-component gases

The correct knowledge of physical theory and models of the mass diffusion for high-temperature multi-component gas flows is of importance for accurately simulating the aerothermodynamic properties of hypersonic vehicles. Mass diffusion mechanism from the phenomenological theory (the Fick’s and the modified Fick’s model) and the kinetic theory (the self-consistent effective binary diffusion model and the Stefan-Maxwell model) is theoretically analyzed to explore the reasons for model differences in application to CFD simulations of Earth reentry flows. Theoretical analysis shows that the other three models rather than the Fick’s model ensure the self-consistent condition. A difference term Ai is proposed to represent the discrepancy of the self-consistent effective binary diffusion model and the modified Fick’s model. Numerical results indicate that whether the mass diffusion model satisfying the self-consistent condition has great influence on surface heat transfer, with the maximum difference up to 18% under the fully-catalytic wall. The aerothermodynamic properties obtained by the modified Fick’s model and the self-consistent effective binary diffusion model are in good agreement, which can be explained by the small Ai values. It is proved that theoretical derivations and numerical results are consistent to conclude that the modified Fick’s model and the self-consistent effective binary diffusion model of rigorous physical fundamental are suitable to describe the mass diffusion process of high-temperature multi-component gases.

Stagnation-point heating and ablation analysis of orbital re-entry experiment

In this study, stagnation-point heating and ablation analysis of the orbital re-entry experiment (OREX) are performed including the air and ablation gas mixture. In gas–gas interactions, the ablation gas is ejected into the shock layer, and the interaction between the air and ablation gases is fully considered. The two-temperature model is employed to describe the thermochemical nonequilibrium flows of the OREX flight conditions. The state-of-art chemical-kinetic parameters of 19-species, including the air and carbon-related ablation gas species, are assessed and utilized to calculate the re-entry flows. In gas–surface interactions, three types of ablation models, the fully equilibrium model, Park model, and surface thermochemistry model of the Zhluktov–Abe and Prata models, are employed to describe the ablation on the surface of carbon–carbon composite CC material of the thermal protection system. For the selected trajectory points of the OREX flight conditions, quasi-one-dimensional thermochemical nonequilibrium flow calculations are carried out, and the results are analyzed in detail. From the calculated results of the re-entry flows, it was found that the production of CO, CO2, and CN is the dominant mechanism of the surface heating on the ablating surface. Heat loss by surface recession is relatively small in OREX flight conditions. The total amount of surface recession due to ablation is approximately 0.22–0.32 mm in the selected range of the OREX flight. Heat loss from surface radiation increases with the surface temperature, and the amount of heat loss is comparable to the amount of surface heating at the trajectory point of 7481.5 s in the OREX flight.

Technique Development for Investigating Axisymmetric Ablation Models in Hypersonic Impulse Facilities

Arcjet facilities have been used extensively for ablation and materials characterization but arcjets do not correctly simulate the hypervelocity flow during reentry. Impulse facilities can simulate aspects of hypervelocity reentry flows, but cannot duplicate the extended-duration thermal environments. To address such deficiencies in short-duration testing environments, a novel plasma technique for preheating axisymmetric samples of heat-shield materials has been developed. The heating technique has been demonstrated using a short-duration blow-down arrangement at the University of Southern Queensland. A 50-mm-diameter graphite disk sample with a uniform thickness of 2 mm was heated from the downstream side with plasma to approximately 2500 K, and then it was exposed to a cold Mach 4.5 flow. The disk was mounted on an experimental probe that was very similar to the European standard probe, normal to the flow. Future applications for the technique are expected to be found in expansion tube facilities that can simulate the true flow energy under reentry conditions.

Modification of chemical-kinetic parameters for 11-air species in re-entry flows

The chemical-kinetic parameters for the two-temperature model for 11-air species are modified to increase the accuracy of the thermochemical nonequilibrium model for high temperature gases in Earth re-entry flows. In the present work, a set of modified chemical-kinetic parameters, for the vibrational relaxation time with high temperature correction, the Arrhenius rate coefficients for dissociation, the vibrational energy removal ratio due to chemical reaction, the equilibrium constant to evaluate the reverse rate coefficients, and collision integrals for high temperature gases are proposed, based on the results of recent state-to-state kinetic calculations and shock-tube experimental data. These modified chemical-kinetic parameters are utilized to reproduce the measured NO, N2, and N2+ VUV and UV spectrum of the high-enthalpy shock-tube experiments performed at the EAST facility, and the existing flight and ground experimental data from BSUV-2, the Mars-entry spacecraft model, RAM C-II, and FIRE-II. The proposed chemical-kinetic parameters more accurately reproduce the measured molecular spectra in the VUV and UV ranges, and the surface heat flux and electron number density on the surface of the re-entry modules, than the previous chemical-kinetic parameters in the Park model, except for the super orbital re-entry speed of the Fire-II experiments. In the Fire-II experimental case, the influence of the modified chemical-kinetic parameters in the present work are less effective because of the fast dissociation of the molecules.

Thermal Environment and Aeroheating Mechanism of Protuberances on Mars Entry Capsule

Mars has only thin atmosphere composed mainly of carbon dioxide that differs significantly from the atmosphere of Earth in terms of characteristics of reentry flows. To connect with the orbiter, the Mars entry capsule is provided with titanium pipes and other units installed on the heat-shield. These units will create significant local interaction flow on the surface of the capsule and cause additional heating on the surface of the shield during the entry of the capsule. With a view to interaction thermal environment issues for the surface of the shield, in this paper, the characteristics of protrusion interaction flow on different location of the shield were studied by means of numerical simulation. Heating mechanisms of protuberances on different location were derived by analyzing characteristic parameters such as local flow velocity, pressure, and Mach number. The results show that the interaction thermal environment of protuberances in the windward area is smaller than that of protuberances in the leeward area, mainly because subsonic flow dominates in the windward area, and the interaction is weak, while in the leeward area, the direction of flow intersects with protuberances to form a boundary layer shear flow, which results in a stronger interaction before the protuberances.

Development of a stagnation streamline model for thermochemical nonequilibrium flow

A stagnation streamline model incorporating quantum-state-resolved chemistry is proposed to study hypersonic nonequilibrium flows along the stagnation streamline. This model is developed by reducing the full Navier–Stokes equations to the stagnation streamline with proper approximations for equation closure. The thermochemical nonequilibrium is described by either the state-to-state approach for detailed analysis or conventional two-temperature models for comparison purpose. The model is validated against various data, and nearly identical results are obtained as compared with those from full field computational fluid dynamics data. In addition, the calculated distributions agree well with the measurement data of a shock tube experiment for the dissociation and vibrational relaxation of O2, including the distributions of species mole fractions and vibrational temperature of the first excited state of O2 molecules. Furthermore, the results with the state-resolved chemistry show that the flow within a shock layer exhibits a strong thermochemical nonequilibrium behavior, which is beyond the capability of commonly used two-temperature models to correctly evaluate the dissociation rate and the associated reaction energy. The present model is also employed to calculate the nonequilibrium re-entry flow along the stagnation streamline for a five-species air mixture as an example to demonstrate the model capability. It is found that both species and internal energy are in a nonequilibrium state, especially the vibrational distributions are strongly deviated from the Boltzmann distribution right behind the bow shock and near the wall surface. The results demonstrate that the proposed stagnation streamline model is very useful to understand thermochemical nonequilibrium phenomena in hypersonic flows.

DSMC modeling of rarefied ionization reactions and applications to hypervelocity spacecraft reentry flows

The DSMC modeling is developed to simulate three-dimensional (3D) rarefied ionization flows and numerically forecast the communication blackout around spacecraft during hypervelocity reentry. A new weighting factor scheme for rare species is introduced, whose key point is to modify the corresponding chemical reaction coefficients involving electrons, meanwhile reproduce the rare species in resultants and preserve/delete common species in reactants according to the weighting factors. The resulting DSMC method is highly efficient in simulating weakly inhomogeneous flows including the Couette shear flow and controlling statistical fluctuation with high resolution. The accurate reliability of the present DSMC modeling is also validated by the comparison with a series of experimental measurements of the Shenzhou reentry capsule tested in a low-density wind tunnel from the HAI of CARDC. The obtained electron number density distribution for the RAM-C II vehicle agrees well with the flight experiment data, while the electron density contours for the Stardust hypervelocity reentry match the reference data completely. In addition, the present 3D DSMC algorithm can capture distribution of the electron, N+ and O+ number densities better than the axis-symmetric DSMC model. The introduction of rare species weighting factor scheme can significantly improve the smoothness of the number density contours of rare species, especially for that of electron in weak ionization case, while it has negligible effect on the macroscopic flow parameters. The ionization characteristics of the Chinese lunar capsule reentry process are numerically analyzed and forecasted in the rarefied transitional flow regime at the flying altitudes between 80 and 97 km, and the simulations predict communication blackout altitudes which are in good agreement with the actual reentry flight data. For the spacecraft reentry with hypervelocity larger than the second cosmic speed, it is forecasted and verified by the present DSMC modeling that ionization reactions will cover the windward capsule surface, leading to reentry communication blackout, and the communication interruption must be considered in the communication design during reentry in rarefied flow regimes.

On the calculation of the electron temperature flowfield in the DSMC studies of ionized re-entry flows

Currently available procedures of electron temperature calculations in studying ionized flows around reentry spacecraft by the direct simulation Monte Carlo (DSMC) method are analyzed. It is shown that the heat conduction of electrons is not taken into account in these procedures. The contributions of various effects to the electron energy balance are calculated by an example of the RAM-C II capsule, and a numerical solution of the electron energy conservation equation is obtained, which refines the electron temperature distribution used in the DSMC computations. A method of coupled calculation of the electron temperature within the framework of the continuum approach and modelling of ionized gas flow by the DSMC method is proposed.

Collision cross sections and nonequilibrium viscosity coefficients of N2 and O2 based on molecular dynamics

This study examines the collision dynamics of atom–atom, atom–molecule, and molecule–molecule interactions for O–O, N–N, O2–O, N2–N, O2–N, N2–O, O2–O2, N2–N2, and N2–O2 systems under thermal nonequilibrium conditions. Investigations are conducted from a molecular perspective using accurate O4, N4, and N2O2 ab initio potential energy surfaces and by performing Molecular Dynamics (MD) simulations. The scattering angle and collision cross sections for these systems are determined, forming the basis for better collision simulations. For molecular interactions, the effect of the vibrational energy on the collision cross section is shown to be significant, which in turn has a profound effect on nonequilibrium flows. In contrast, the effect of the rotational energy of the molecule is shown to have a negligible effect on the cross section. These MD-based cross sections provide a theoretically sound alternative to the existing collision models, which only consider the relative translational energy. The collision cross sections reported herein are used to calculate various transport properties, such as the viscosity coefficient, heat conductivity, and diffusion coefficients. The effect of internal energy on the collision cross sections reflects the dependence of these transport properties on the nonequilibrium degree. The Chapman–Enskog formulation is modified to calculate the transport properties as a function of the trans-rotational and vibrational temperatures, resulting in a two-temperature nonequilibrium model. The reported work is important for studying highly nonequilibrium flows, particularly hypersonic re-entry flows, using either particle methods or techniques based on the conservation laws.