Aerothermal Environment Research Articles

Fabrication of Sodium/Inconel 718 Heat Pipes for Combustor Cooling

Cooling channel involved heat pipes, which synergistically combined the high conductivity of heat pipe and thermal efficiency of regenerative cooling, were considered for combustor cooling. In this study, cooling channels were fabricated in sodium/Inconel 718 heat pipes in China Academy of Aerospace Aerodynamics (CAAA). And their startup properties and thermal response were investigated systematically. It was found that the frozen startup limit of fabricated heat pipes were 2.60 and 2.76, satisfying the criterion for frozen startup. During startup tests, the sodium/Inconel 718 heat pipes startup successfully, displaying uniform temperatures of about 900 K and 830 K, respectively. Moreover, sodium/Inconel 718 heat pipes decreased operating temperatures and alleviated thermal stresses in combustor sidewalls. Under simulated Ma6 aerothermal environment, the operating temperature of sodium/Inconel 718 heat pipe was 1007 K, being 224 K lower than that of regenerative cooling model. Therefore, it was concluded that heat pipes could increase the thermal margin of regenerative cooling system, being suitable for applications in combustor cooling.

Microstructure, thermophysical properties, and ablation resistance of C/HfC-ZrC-SiC composites

Carbon/carbon composites modified by HfC-ZrC-SiC were fabricated by reactive melt infiltration with the aim of improving their ablation resistance for application in aerothermal environments. Their microstructure, thermophysical and ablation properties were investigated. Results show that the thermal diffusivity decreases with increasing temperature for all composites. The thermal conductivity of the C/HfC-ZrC-SiC composites decreasing with increasing HfC molar fraction is related to decreased grain size and increased porosity, which impede phonon interaction and increase the phonon scattering. High HfC content effectively improves the oxidation and ablation resistance of the composites. C/HfC-ZrC-SiC composites containing 8.8 mol.% HfC exhibited the best ablation resistance owing to a compact and continuous HfO2-ZrO2 mixed layer that formed on the ablated surface.

Multidimensional material response simulations of a full-scale tiled ablative heatshield

The Mars Science Laboratory (MSL) was protected during Mars atmospheric entry by a 4.5 meter diameter heatshield, which was constructed by assembling 113 thermal tiles made of NASA's flagship porous ablative material, Phenolic Impregnated Carbon Ablator (PICA). Analysis and certification of the tiles thickness were based on a one-dimensional model of the PICA response to the entry aerothermal environment. This work provides a detailed three-dimensional heat and mass transfer analysis of the full-scale MSL tiled heatshield. One-dimensional and three-dimensional material response models are compared at different locations of the heatshield. The three-dimensional analysis is made possible by the use of the Porous material Analysis Toolbox based on OpenFOAM (PATO) to simulate the material response. PATO solves the conservation equations of solid mass, gas mass, gas momentum and total energy, using a volume-averaged formulation that includes production of gases from the decomposition of polymeric matrix. Boundary conditions at the heatshield forebody surface were interpolated in time and space from the aerothermal environment computed with the Data Parallel Line Relaxation (DPLR) code at discrete points of the MSL trajectory. A mesh consisting of two million cells was created in Pointwise, and the material response was performed using 840 processors on NASA's Pleiades supercomputer. The present work constitutes the first demonstration of a three-dimensional material response simulation of a full-scale ablative heatshield with tiled interfaces. It is found that three-dimensional effects are pronounced at the heatshield outer flank, where maximum heating and heat loads occur for laminar flows.

Thermochemical ablation modeling forward uncertainty analysis—Part I: Numerical methods and effect of model parameters

Next generation spacecraft will bring back heavier payloads from explored planets. Advance in the modeling of the thermo-chemical ablation of carbon-based thermal protection system materials is fundamental to improve the design capabilities of these vehicles. Computational fluid dynamic approaches are extensively used to model the gas-surface interaction phenomena over ablative materials. The advantage of such kind of approaches is the accurate description of the aerothermal environment obtained through the full resolution of the mechanical, thermal, and chemical boundary layers that develop over an ablative surface when exposed to a high-enthalpy flow. This paper is devoted to the assessment of the uncertainties of such kind of thermo-chemical ablation model and to study their effect on the model final outcomes. A sphere of non-pyrolyzing carbon-based material, exposed to conditions similar to those of a typical plasma wind tunnel test, is the selected test case for the analysis. Two forward non-intrusive uncertainty quantification techniques are used to analyze the influence of the defined set of uncertain parameters on the estimate of steady-state mass blowing flux and surface temperature. Our results show that for the selected conditions, and uncertainty ranges, the surface nitridation reaction probability has the strongest impact on the model outcomes.

热环境条件下红外窗口气动光学传输效应实验研究

热环境条件下红外窗口气动光学传输效应实验研究

Numerical analysis of Hall effect on the performance of magnetohydrodynamic heat shield system based on nonequilibrium Hall parameter model

Magnetohydrodynamic (MHD) heat shield system, a novel thermal protection technique in the hypersonic field, has been paid much attention in recent years. In the real flight condition, not only the Lorentz force but also the Hall electric field is induced by the interaction between ionized air post shock and magnetic field. In order to analyze the action mechanisms of the Hall effect, numerical methods of coupling thermochemical nonequilibrium flow field with externally applied magnetic field as well as the induced electric field are constructed and validated. Based on the nonequilibrium model of Hall parameter, numerical simulations of the MHD heat shield system is conducted under two different magnetic induction strengths (B0=0.2T, 0.5T) on a reentry capsule forebody. Results show that, the Hall effect is the same under the two magnetic induction strengths when the wall is assumed to be conductive. For this case, with the Hall effect taken into account, the Lorentz force counter stream diminishes a lot and the circumferential component dominates, resulting that the heat flux and shock-off distance approach the case without MHD control. However, for the insulating wall, the Hall effect acts in different ways under these two magnetic induction strengths. For this case, with the Hall effect taken into account, the performance of MHD heat shield system approaches the case neglecting the Hall effect when B0 equals 0.2T. Such performance becomes worse when B0 equals 0.5T and the aerothermal environment on the capsule shoulder is even worse than the case without MHD control.

Optics Temperature-Dependent Nonuniformity Correction Via <inline-formula> <tex-math notation="LaTeX"> $\ell _{0}$</tex-math> </inline-formula>-Regularized Prior for Airborne Infrared Imaging Systems

In this paper, we propose a new $\ell _{0}$ -regularized approach to remove the temperature-dependent nonuniformity effects induced by the infrared (IR) imaging optics in an aerothermal environment. The $\ell _{0}$ image prior is inspired by observing distinct characteristics of IR images with small targets. Based on this effective prior, we present a variational framework where we optimize an energy functional to estimate the optics-related fixed pattern noise (FPN) and the latent image. A computationally efficient numerical algorithm based on half-quadratic regularization is used to solve the optimization problem. The proposed method is fundamentally different from the existing nonuniformity correction techniques developed for infrared focal plane arrays and simultaneously suppresses the optics-related FPN and random noise. Both quantitative and qualitative comparisons to specialized state-of-art algorithms demonstrate its superiority.

Stability analysis and vibration characteristics of an axially moving plate in aero-thermal environment

The stability and vibration characteristics of an axially moving plate in an aero-thermal environment subjected to transverse excitation are investigated. In the modeling of the equation of motion, the influences of the in-plane thermal load and the perturbation aerodynamic pressure on the transverse bending deflection of the axially moving plate are taken into account. The governing equation of the plate in aero-thermal environment is established applying the Hamilton’s principle based on the von Karman nonlinear plate theory and linear potential flow theory. For the linear equation, the natural frequencies of the moving plate are analyzed by solving the generalized eigenvalue problem. The critical parameters of the moving velocity, flow velocity and temperature change for the divergence of the plate are obtained. For the nonlinear equation, the displacement time responses of the plate in different stability states subjected to transverse excitation are analyzed by numerical simulations. From the study, it can be seen that with the moving velocity, flow velocity and temperature change increasing, the fundamental natural frequency of the plate decreases. When the fundamental natural frequency decreases to 0, the plate is in a divergence type of instability. The critical moving velocity decreases with increasing flow velocities and temperature changes. The vibration amplitude of the plate in divergence state is larger than that in the stable state. The vibration amplitude increases with the flow velocity and temperature change increasing, which illustrates that the aero-thermal environment has significant effects on the stability and vibration properties of the axially moving plate.

Fluid–Thermal Interaction Investigation of Spiked Blunt Bodies at Hypersonic Flight Condition

A spiked blunt body flying at hypersonic speed could result in significant thermal interactions between the external flowfield and the internal structure, which have not received enough attention in previous investigations. In the present study, time-dependent, loosely coupled fluid–thermal interaction simulations are conducted on a blunt body equipped with a spike with variable length at a , flight condition. The results show that, in the range of time considered in the study, a blunt body with a longer spike yields a better dynamic thermal performance, that is, a lower peak temperature on the blunt forebody surface and a slower temperature rise both on the forebody surface and inside the structure can be obtained. In addition, various levels of drag coefficient drop can be observed during the coupling process. A maximum decrease of 5.7% is achieved among the investigated models. The study reveals the necessity of the coupled fluid–thermal analysis to accurately predict the aerothermal environment of the spiked blunt body and to capture the variation of aerodynamic force, which is critical for structural design and flight control system, respectively.

Aero-thermal analysis of lifting body configurations in hypersonic flow

The aero-thermal analysis of a hypersonic vehicle is of fundamental interest for designing its thermal protection system. The aero-thermal environment predictions over several critical regions of the hypothesized lifting body vehicle, including the stagnation region of the nose-cap, cylindrically swept leading edges, fuselage-upper, and fuselage-lower surfaces, are discussed. The drag (Λ=70°) and temperature (Λ=80°) minimized sweepback angles are considered in the configuration design of the two hypothesized lifting body shape hypersonic vehicles. The main aim of the present study is to analyze and compare the aero-thermal characteristics of these two lifting body configurations at same heat capacity. Accordingly, a Computational Fluid Dynamics simulation has been carried out at Mach number (M∞=7), H=35km altitude with zero Angle of Attack. Finally, the material selection for thermal protection system based on these predictions and current methodology is described.

Modeling of Heat Transfer Attenuation by Ablative Gases During the Stardust Reentry

Modern space vehicles designed for planetary exploration use ablative materials to protect the payload against the high heating environment experienced during reentry. To properly model and predict the aerothermal environment of the vehicle, it is imperative to account for the gases produced by ablation processes. The present study aims to examine the effects of the blowing of ablation gas in the outer flow field. Using six points on the Stardust entry trajectory at the beginning of the continuum regime, from 81 to 69 km, the various components of the heat flux are compared to air-only solutions. Although an additional component of the heat flux is introduced by mass diffusion, this additional term is mainly balanced by the fact that the translational–rotational component of the heat flux, the main contributor, is greatly reduced. Although a displacement of the shock is observed, it is believed that the most prominent effects are caused by a modification of the chemical composition of the boundary layer, which reduces the gas-phase thermal conductivity.

On-ground casualty risk reduction by structural design for demise

In recent years, awareness concerning the on-ground risk posed by un-controlled re-entering space systems has increased. On average over the past decade, an object with mass above 800kg re-enters every week from which only a few, e.g. ESA’s GOCE in 2013 and NASA’s UARS in 2011, appeared prominent in international media. Space agencies and nations have discussed requirements to limit the on-ground risk for future missions. To meet the requirements, the amount of debris falling back on Earth has to be limited in number, mass and size. Design for demise (D4D) refers to all measures taken in the design of a space object to increase the potential for demise of the object and its components during re-entry. SCARAB (Spacecraft Atmospheric Re-entry and Break-Up) is ESA’s high-fidelity tool which analyses the thermal and structural effects of atmospheric re-entry on spacecraft with a finite-element approach. For this study, a model of a representative satellite is developed in SCARAB to serve as test-bed for D4D analyses on a structural level. The model is used as starting point for different D4D approaches based on increasing the exposure of the satellite components to the aero-thermal environment, as a way to speed up the demise. Statistical bootstrapping is applied to the resulting on-ground fragment lists in order to compare the different re-entry scenarios and to determine the uncertainties of the results. Moreover, the bootstrap results can be used to analyse the casualty risk estimator from a theoretical point of view. The risk reductions for the analysed D4D techniques are presented with respect to the reference scenario for the modelled representative satellite.

Evaluation of homogenized effective properties for FGM panels in aero-thermal environments

Based on homogenization methods, effective properties of functionally graded (FG) panels are investigated for the two types of models such as the power law (P-) and the sigmoid (S-) FGM functions. For heterogeneous micromechanical model, the effective material properties of the structures are obtained using Voigt rule (VR) and Mori–Tanaka method (MTM). The governing equations are derived by the first-order shear deformation plate theory with von Karman strain–displacement relationships. And then, the effective properties of the FGMs are evaluated by VR and MTM in detail. The results show that the characteristics of FGM structures are determined dominantly according to the constituents. In order to investigate the macroscopic behaviors, Newmark time integration method is applied to analyze the aero-thermoelastic behaviors of the models based on the first-order piston theory. Also, nonlinear dynamic behaviors of the panels are studied under supersonic aerodynamic force.

Reconstruction of Aerothermal Environment and Heat Shield Response of Mars Science Laboratory

An initial assessment and reconstruction of the Mars Science Laboratory entry aerothermal environment and thermal protection system response is performed using the onboard instrumentation suite called the Mars Science Laboratory entry, descent, and landing instrumentation. The analysis is performed using the current best-estimated trajectory. The Mars Science Laboratory Entry, Descent, and Landing Instrumentation suite in part provides in-depth temperature measurements at seven locations on the heat shield. The temperature data show the occurrence of boundary-layer transition to turbulence on the leeside forebody of the entry vehicle. The data also suggest that the thermal protection system recession is lower than nominal model predictions using diffusion limited surface oxidation. The model predictions of temperatures show an underprediction in the stagnation and apex regions and an overprediction in the leeside region. An estimate of time-varying aeroheating using an inverse reconstruction technique is also presented. The reconstructed aeroheating is sensitive to the choice of a recession model. A few areas of excess margins in aerothermal environments and thermal protection system sizing are identified for reevaluation in future designs.

Two-Dimensional Ablation and Thermal Response Analyses for Mars Science Laboratory Heat Shield

This paper examines transient simulations performed to predict in-depth thermal response and surface recession of the proposed heat shield material for the Mars Science Laboratory entry capsule, that is, phenolic impregnated carbon ablator. The finite volume material response code used in this paper solves the time-dependent governing equations, including energy conservation and a three-component decomposition model, with a surface energy-balance condition and a moving grid system to predict shape change due to surface recession. The predicted in-depth thermal response of heat shield material generally agrees well with the thermocouple data under various arcjet conditions. Also, two-dimensional computations using aerothermal environment for Mars entry (derived from a proposed three-sigma trajectory) are performed around the heat shield shoulder region, where high heating occurs as the result of angle of attack. Parametric studies are conducted to examine the effects of carbon-fiber orientation, material properties, and surface recession on heat shield bondline temperature history. It is proved that the fiber orientation configuration of the baseline heat shield has the lowest maximum bondline temperature.

Mars Science Laboratory Entry Atmospheric Data System Trajectory and Atmosphere Reconstruction

On 5 August 2012, the Mars Science Laboratory entry vehicle successfully entered Mars’s atmosphere and landed the Curiosity rover on its surface. The entry vehicle carried a system of heat shield instrumentation that measured the aerodynamic and aerothermal environment during entry. Included in these sensors were seven pressure transducers linked to ports across the entry vehicle forebody that recorded the local surface pressures during entry. These measured surface pressures were used to generate estimates of atmospheric quantities based on computationally modeled surface pressure distributions. Angle of attack, angle of sideslip, dynamic pressure, Mach number, atmospheric conditions, and vehicle aerodynamics were reconstructed from the measured pressures. Three data reduction algorithms using subsets of the available data were used to assess data quality and interpret the entry performance. A Kalman filter algorithm was used to combine all available data for a final, complete reconstruction. The reconstruction results indicate that the data quality was good, and aerodynamic performance was within the uncertainties of preflight models. The three independent reconstructions are in overall good agreement, with several small anomalies that were reconciled using reasonable corrections within known sources of error. The combined Kalman filter reconstruction identified winds that are reasonable compared with preflight atmospheric models.

Feasibility of Guided Entry for a Crewed Lifting Body Without Angle-of-Attack Control

The feasibility of flying a crewed lifting body, such as the HL-20, during entry from low Earth orbit without the steady-state body flap deflections required for angle-of-attack control was evaluated. This entry strategy mitigates the severity of the aerothermal environment on the vehicle’s body flaps and reserves control power for transient steering maneuvers. A real-time numeric predictor–corrector entry guidance algorithm was developed to accommodate the range of vehicle trim angle-of-attack profiles possible in the absence of angle-of-attack control. Results show that it is feasible to steer the vehicle from low Earth orbit to a desired target with real-time guidance while satisfying a reasonable suite of trajectory constraints, including limits on peak heat rate, peak sensed deceleration, and integrated heat load. Uncertainty analyses confirm this result and show that the vehicle maintains significant performance robustness to expected day-of-flight uncertainties. Additionally, parametric scans over mission design parameters of interest indicate that a high level of flexibility is available for the low-Earth-orbit return mission. Together, these results indicate that the proposed entry strategy is feasible: Crewed lifting bodies may be effectively flown without steady-state body flap deflections.

Analysis and modeling of aerothermal radiation based on experimental data

The effect of aerothermal radiation on the performance of an airborne infrared imaging system is becoming a crucial issue in the development of electro-optic systems. In this paper the high-temperature radiation characteristics of infrared window are studied in depth, and a multi-scale analysis model of aerothermal radiation degraded images is presented. The presented model adopted a least-squares fitting method to simulate the salient characteristics of the aerothermal radiation degraded images. The aerothermal radiation degradation characteristic database containing characterizing parameters of degraded images was created. Through optimization selection, the characterizing parameters were used to correct the degraded images in the real aerothermal environment. The experimental results show that the model is effective in the aerothermal radiation degraded image correction, and improves the quality of the aerothermal radiation degraded image.

Study of Ablation-Flowfield Coupling Relevant to the Orion Heatshield

The coupled interaction between an ablating surface and the surrounding aerothermal environment is studied. An equilibrium ablation model is coupled to the LAURA flowfield solver, which allows the char ablation rate (m-dot(sub c)) to be computed as part of the flowfield solution. The wall temperature (T(sub w)) and pyrolysis ablation rate (m-dot(sub g)) may be specified by the user, obtained from the steady-state ablation approximation, or computed from a a material response code. A 32 species thermochemical nonequilibrium flowfield model is applied, which permits the treatment of C, H, O, N, and Si containing species. Coupled ablation cases relevant to the Orion heatshield are studied. These consist of diffusion-limited oxidation cases with Avcoat as the ablation material. The m-dot(sub c) values predicted from the developed coupled ablation analysis were compared with those obtained from a typical uncoupled ablation analysis. The coupled results were found to be as much as 50% greater than the uncoupled values. This is shown to be a result of the cumulative effect of the two fundamental approximations inherent in the uncoupled analysis.

High performance modeling of atmospheric re-entry vehicles

Re-entry vehicles designed for space exploration are usually equipped with thermal protection systems made of ablative material. In order to properly model and predict the aerothermal environment of the vehicle, it is imperative to account for the gases produced by ablation processes. In the case of charring ablators, where an inner resin is pyrolyzed at a relatively low temperature, the composition of the gas expelled into the boundary layer is complex and may lead to thermal chemical reactions that cannot be captured with simple flow chemistry models. In order to obtain better predictions, an appropriate gas flow chemistry model needs to be included in the CFD calculations. Using a recently developed chemistry model for ablating carbon-phenolic-in-air species, a CFD calculation of the Stardust re-entry at 71 km is presented. The code used for that purpose has been designed to take advantage of the nature of the problem and therefore remains very efficient when a high number of chemical species are involved. The CFD result demonstrates the need for such chemistry model when modeling the flow field around an ablative material. Modeling of the nonequilibrium radiation spectra is also presented, and compared to the experimental data obtained during Stardust re-entry by the Echelle instrument. The predicted emission from the CN lines compares quite well with the experimental results, demonstrating the validity of the current approach.