Aerothermodynamic Environment Research Articles

A fully coupled multiscale phase-change model at the porous interface for transpiration cooling: coupling dynamics pore-scale networks to continuum-scale free flow

A fully coupled multiscale phase-change model at the porous interface for transpiration cooling: coupling dynamics pore-scale networks to continuum-scale free flow

A Multifidelity Framework for Aerothermodynamic Modeling and Simulation of Destructive Atmospheric Entry

The aerothermodynamic environment surrounding bodies undergoing destructive atmospheric entry is characterized by multiple interacting shock waves generated by the presence of several fragments moving close to one another at high speed. Shock interaction and impingement produce highly localized surface loads, influencing the overall dynamics of the fragment, its temperature, and in turn its demise. To account for the relevant physics during the reentry process of fragmenting bodies, a multifidelity framework is presented that allows for selecting the most appropriate level of fidelity, between high-fidelity solutions or low-fidelity aerothermodynamic metamodels. A physics-informed approach is implemented to automatically choose between the two fidelity levels on the basis of the intersection of shock wave envelopes. The complete framework is validated on relevant test cases from the literature and its performance is demonstrated for destructive atmospheric reentry cases, such as the case of the Attitude Vernier Upper Module and the Automated Transfer Vehicle from the European Space Agency.

Multiscale coupling simulation of surface catalytic effect on hypersonic aerothermodynamic environment

Multiscale coupling simulation of surface catalytic effect on hypersonic aerothermodynamic environment

Analysis of the integrated performance of hybrid fiber-reinforced polymer composite used for thermal protection based on a dual-scale ablation model

Analysis of the integrated performance of hybrid fiber-reinforced polymer composite used for thermal protection based on a dual-scale ablation model

Influence of vehicle length on the aerothermodynamic environment of the Hyperloop

Influence of vehicle length on the aerothermodynamic environment of the Hyperloop

Nonlinear finite element algorithm for solving fully coupled thermomechanical problems under strong aerothermodynamic environment

Nonlinear finite element algorithm for solving fully coupled thermomechanical problems under strong aerothermodynamic environment

Numerical analysis of deformation and failure mechanism of metal truss structure of spacecraft under re-entry aerothermodynamic environment

A coupled aerodynamic-thermoelastic analysis framework is presented to simulate the deformation and failure mechanism of metal truss structure for spacecraft under re-entry aerothermodynamic environment. The three-dimensional dynamic thermo-mechanical coupling finite element algorithm is developed by computable modeling of thermoelasticity and heat conduction equations to solve the thermo-mechanical coupling response of spacecraft structure. The aerodynamic boundary conditions such as temperature, heat flux and pressure distribution are computed and provided by the gas-kinetic unified algorithm (GKUA) based on the Boltzmann model equation. The linear interpolation method is designed and employed to couple the boundary conditions between the external aerodynamic flow field and the solid structure interface. The present thermo-mechanical coupling algorithm is verified on one-dimensional and two-dimensional transient heat flow problems with analytical solutions, and good consistency is achieved. The reliability of the GKUA for solving the boundary conditions of the external flow field and the dynamic thermo-mechanical coupling algorithm for the thermal response behavior of the internal structure material is verified by the consistency calculation of the internal and external temperature distribution. Then, the aerodynamic-thermoelastic analysis framework is further applied to study the deformation behavior of structural thermo-mechanical responses and failure mechanism of a hollow sphere under re-entry aerothermodynamic environment with emphasis placed on the influence of flight speeds, rarefied effects, algorithm schemes and damping effects . The numerical results obtained by the presented algorithm including the internal temperature, thermal stress and deformation distribution law are compared and analyzed. Finally, the deformation and failure mechanism of spacecraft structure during re-entry process are solved and analyzed by the present algorithm framework. The displacement, temperature and stress distribution contours of the structure are presented to make a preliminary understanding and evaluation on the process of spacecraft deformation and failure disintegration. It is indicated that the present coupled aerodynamic-thermoelastic analysis framework can provide a reliable, applicable and efficient platform for the thermo-mechanical coupling responses and failure mechanism of re-entry spacecraft in the end of life.

Comprehensive performance of multifunctional lightweight composite reinforced with integrated preform for thermal protection system exposed to extreme environment

The novel multifunctional lightweight composites (MLCs) reinforced with integrated preform were prepared for thermal protection system (TPS) of hypersonic vehicles in extreme aerothermodynamic environment. The high-temperature behavior including thermal physical properties decomposition performance, and the ablation resistance were tested by the oxy-acetylene torch system, as well as the micro-morphologies of MLCs from non-destructive testing were investigated. An integration factor was proposed to evaluate the comprehensive performance of MLCs. The results show that MLCs had high structural efficiency without delamination and defelting, excellent thermal insulation performance, and superior ablation resistance ability. The comprehensive performance of MLCs had obvious advantages to polymer-matrix composites (PMCs). The novel MLC reinforced with integrated preform can help for the development of thermal protection materials of hypersonic aircrafts.

Joule heated hot wall models in hypersonic wind tunnel facilities for energy deposition studies

Joule heating is a useful experimental technique for ground based hypersonic testing as it allows replication of the governing wall-to-freestream temperature ratio and thus better simulation of the aerothermodynamic environment of high-speed flight vehicles. This paper presents design considerations for using this hot-wall method for hypersonic wind tunnel models. A numerical analysis of material choice and thickness for creating rapid uniform high-temperature model conditions is presented. The influence of the method and location of energisation and structural requirements for the integration of joule heating in a hypersonic wind tunnel model is also examined numerically. Experimental results of a bench-top study incorporating the design elements for joule heating for studies analogues to thermal energy deposition are also presented. To replicate thermal energy deposition, graphite plates with a 1.0 or 2.5 mm thickness were heated via joule heating to a base level of either 30 W or 100 W. After 120 s of heating the plates were then exposed to a thermal pulse of either 550, 1200 or 1300 W. The thermal-structural response of the plate was experimentally recorded with a measurement system incorporating infrared thermal cameras, single and two-colour pyrometers, a load cell and a 520 nm light Digital Image Correlation system suitable for use on hot wall models. Results from this bench-top study demonstrate a proof-of-concept for using joule heating for analogue thermal deposition studies. Temperature rises of Δ T on the order of 500 - 1000 K were achieved on models with base temperatures of 450 - 600 K achieved with this technique. Plate deflections as high as 3.2 - 4.1 mm were recorded for 1.0 mm plates with both stress and strain of the heated models resolved experimentally for base heating and pulsed energy conditions.

Competing effects of surface catalysis and ablation in hypersonic reentry aerothermodynamic environment

Under hypersonic flow conditions, the complicated gas-graphene interactions including surface catalysis and surface ablation would occur concurrently and intervene together with the thermodynamic response induced by spacecraft reentry. In this work, the competing effects of surface heterogeneous catalytic recombination and ablation characteristics at elevated temperatures are investigated using the Reactive Molecular Dynamics (RMD) simulation method. A Gas-Surface Interaction (GSI) model is established to simulate the collisions of hyper-enthalpy atomic oxygen on graphene films in the temperature range of 500–2500 K. A critical temperature T c around 900 K is identified to distinguish the graphene responses into two parts: at T < T c , the heterogeneous surface catalysis dominates, while the surface ablation plays a leading role at T > T c . Contradicting to the traditional Arrhenius expression that the recombination coefficient increases with the increase of surface temperature, the value is found to be relatively uniform at T < T c but declines sharply as the surface temperature increases further due to the competing ablation effect. The occurrence of surface ablation decreases the amounts of active sites on the graphene surface for oxygen adsorption, leading to reduced recombination coefficient from both Langmuir-Hinshelwood (L-H) and Eley-Rideal (E-R) mechanisms. It suggests that the traditional Computational Fluid Dynamics (CFD) simulation method, which relies on the Arrhenius-type catalysis model, would result in large discrepancies in predicting aerodynamic heat for carbon-based materials during reentry into strong aerodynamic thermal environment.

Assessment of high enthalpy flow conditions for re-entry aerothermodynamics in the plasma wind tunnel facilities at IRS

This article presents the full operational experimental capabilities of the plasma wind tunnel facilities at the Institute of Space Systems at the University of Stuttgart. The simulation of the aerothermodynamic environment experienced by vehicles entering the atmosphere of Earth is attempted using three different facilities. Utilizing the three different facilities, the recent improvements enable a unique range of flow conditions in relation to other known facilities. Recent performance optimisations are highlighted in this article. Based on the experimental conditions demonstrated a corresponding flight scenario is derived using a ground-to-flight extrapolation approach based on local mass-specific enthalpy, total pressure and boundary layer edge velocity gradient. This shows that the three facilities cover the challenging parts of the aerothermodynamics along the entry trajectory from Low Earth Orbit. Furthermore, the more challenging conditions arising during interplanetary return at altitudes above 70 km are as well covered.

Design of ablation resistant/heat insulation/lightweight integrated thermal protection material for extreme aerothermodynamic environment

AbstractIn order to meet the requirement of the more extreme and severe aerothermodynamic environment faced by the future spacecrafts for high‐speed Earth return missions, novel ablation resistant/heat insulation/lightweight integrated resin matrix composites are designed for the heatshields for these vehicles. Proposing the concept of security zone and heat insulation zone into the novel designs, not only the ablation resistant and heat insulation abilities can be improved, but also the weight of the entire material can be effectively reduced. Simultaneously, a computer modeling is established to evaluate the comprehensive performance of integrated high temperature materials. A variety of material process design parameters is introduced into the model while taking the respective properties of fiber weave and matrix into consideration. Computer codes are written for solving the model based on the numerical method for solving nonlinear equations. Under two typical extreme aerothermodynamic environments faced by future aircrafts, the effects of material process design parameters on the comprehensive performance are analyzed, which includes ablation resistant ability, heat insulation capacity, and lightweight level. Finally, the optimal design scheme of the new integrated resin matrix composite under the specific environment is determined. This work helps to the on‐demand design of new thermal protection materials for extreme entry or reentry environments.

Coupled investigation on drag reduction and thermal protection mechanism of a double-cone missile by the combined spike and multi-jet

Vehicles flying at supersonic/hypersonic speeds would encounter severe aerodynamic and aerothermodynamic environment, and hypersonic vehicles would produce high pressure and heat flux loads. In this work, the fluid-thermal coupling analysis is performed on the double-cone missile. The spike, as well as the combined spike and multi-jet strategy are applied to the blunt head of the missile for drag reduction and thermal protection. The impacts of the jet pressure ratio (PR) and length-to-diameter ratio (L/D) on the flowfield and the aerodynamic characteristics are investigated numerically. The obtained results reveal that the heat flux and static pressure can be reduced significantly by increasing the jet pressure ratio, and the maximum peak pressure and heat flux decrease up to 58.21% and 46.43% respectively. The flow field structure highly depends on the length of the spike, and the total drag coefficient is not monotonically reduced with the increase of the length of the spike. Moreover, it is noted that the aerodynamic heating level and the drag value are both related to the reattachment angle of the reattachment shock wave.

A model for thermal protection ablative material with local thermal non-equilibrium and thermal radiation mechanisms

Novel resin-based ablative materials are manufactured for the Thermal Protection System (TPS) of hypersonic aircraft in recent years to adapt more extreme aerothermodynamic environment. The establishment of theoretical model and simulation method for accurately predicting material's responses can guide design of TPS and significantly reduce Research and Development (R&D) cycle and cost. In order to solve the problem of inaccurate material response calculated by the heritage models, which is caused by the new mechanisms involved in new materials. On the basis of heritage physical and chemical mechanisms for resin-based ablative materials, this study further introduces the mechanisms of the local thermal non-equilibrium and thermal radiation, and a developed thermal-fluid-ablation coupled theoretical model is established. The coupled model is numerically solved by writing computer codes. After completing the verification of the model and simulation codes, the influences of the new mechanisms on the performance of ablative material are analyzed. The results show that the effect of local thermal non-equilibrium and thermal radiation cannot be ignored for current ablative materials, in which the key factors are the pore size and porosity. • Local thermal non-equilibrium and thermal radiation are considered for ablators. • A thermal-fluid-ablation coupled theoretical model is developed with new mechanisms. • The effect of local thermal non-equilibrium and thermal radiation cannot be ignored. • The model and method can guide novel material design and reduce R&D cycle and cost.

Coupled investigation on drag and heat flux reduction over spiked waverider with a wide-speed range

When flying at supersonic/hypersonic speeds, the vehicles constantly suffer from extremely complex aerodynamic and aerothermodynamic environment. In order to investigate the drag reduction and thermal protection characteristics of hypersonic vehicles, the spike, as well as the combined spike and opposing jet strategy, are applied to the waverider. The effects of different strategies and angle of attack on the flow field structure and the aerodynamic characteristics of a waverider are studied, and the fluid-thermal interaction between the three-dimensional flow field and the internal structure is explored. The angle of attack has significant influences on the lift and drag coefficient values, and increasing the angle of attack can further increase the lift and drag coefficient values, but result in the decrease of the lift-to-drag ratio. The coupled analysis results indicate that a prominent reduction in the drag and aeroheating is achieved when the combined spike and opposing jet strategy is added to the waverider, and the flow field structure changes remarkably during the coupling process.

Deriving a Control-Oriented Model for an Axisymmetric Vehicle With the Power-Law Revolution Nose

The purpose of this paper is to construct a new general vehicle model as an open fundamental material for the guidance and control research. In this study, parameterized configuration, aerodynamics calculation, control-oriented modeling, stability analysis, and nominal trajectory design are performed for the general vehicle model. First, the aerodynamic configuration is parameterized as an axisymmetric body with a power-law revolution nose. Then, an engineering method considering inviscid flow, base drag and skin friction is used for the aerodynamics calculation, and a control-oriented fitting model of longitudinal aerodynamics is established based on the analysis of the correlation between aerodynamic force and the parameters of Mach number, attack angle, elevator deflection and height. Next, the aerothermodynamic environment prediction of power-law revolution axisymmetric hypersonic vehicle (PRAHV) is discussed, and the nose heating rate formula of PRAHV is established. The stability analysis and nominal trajectory design of PRAHV is performed based on the fitting model and the heating rate formula. The stability analysis shows that both the static stability and dynamic stability of the vehicle are unstable. The nominal trajectory of unpowered longitudinal maneuvering is achieved by the hp-adaptive pseudospectral method, which demonstrated that the availability of the control-oriented model established in this paper. In conclusion, this work provides a fundamental object for further study of vehicle guidance, control, and evaluation.

Evaluation method and key factor analysis for thermal protection performance of multifunctional integrated ablative materials

AbstractThe thermal protection materials of hypersonic aircrafts in near space are subjected to extremely serious aerothermodynamic environment under long heating time, higher enthalpy, larger heat accumulation, and so forth. Proposing new concepts of materials and accurate evaluation of their performance take the dominant roles in developing high‐efficiency thermal protection materials under this severe environment. By introducing the concepts of gradient and fiber hybridization, this study proposes a multifunctional integrated ablative material scheme to replace heritage ones in order to improve the heat insulation, ablation resistant, and weight reduction capacities. Meanwhile, a thermal‐fluid‐ablative coupling mathematical model taking the manufacturing processing parameters reflecting the gradient and fiber hybridization into consideration is established. Furthermore, the sensitivity analysis for the manufacturing processing parameters is carried out. A simulation tool is developed based on writing FORTRAN codes to solve the multi‐field coupling model numerically. The results that the model and code are effective in analyzing the material's performance and the key manufacturing processing parameters influencing on the thermal protection performance of multifunctional integrated ablative material are pointed out. This work helps to design new thermal protection materials for hypersonic vehicles under extreme aerothermodynamic environments.Highlights New scheme of multifunctional integrated ablative material is proposed. Developed multi‐field coupling thermal protection model for new material is established. A bridge between material responses and manufacture processing parameters is built. Key factors influencing on material performance are pointed out by sensitivity analysis.

Influence factors on the multi-field coupling performances of charring ablators on the basis of a mesoscopic ablation model

• A mesoscopic ablation model is developed to solve coupling performances. • Gasification coefficient almost has no influence on mechanical performance. • Adding fiber can improve thermal and mechanical performances of ablator. • Effective elastic modulus can be predicted by a temperature dependent function. The prediction of the multi-field (thermal-fluid-chemical-mechanical) coupling performances of charring ablators used in integral thermal protection system (ITPS) for reentry vehicles is subject to the aerothermodynamic environment. A mesoscopic ablation model is built in this manuscript, which takes the mass, momentum, energy conservations and the mesoscopic mechanics model for the solid and the gaseous phases produced from the ablation process into consideration. Based on this model, firstly, the thermal responses and the effective modulus for existing materials are calculated. The simulation results are compared with the experimental data to validate our developed mesoscopic ablation model. Then, the heat and mass transfer of each phase in the ablator, the fluid performance of the pyrolysis gases, the chemical responses of the pyrolysis reactions, as well as the effective mechanical properties under ultra-temperature for charring ablators with different influence factors (the gasification coefficient and the fiber’s volume fraction) are simulated and analyzed. The numerical results indicate that the variation of the fiber’s volume fraction has significant influence on the coupled ablative performance. The effective elastic modulus of the ablator can be predicted by a temperature dependent linear function. The pyrolysis reactions proceed completely with producing large amounts of pyrolysis gases in the ablator with a low effective elastic modulus.

Effects of recess depth and opening diameter on pressure measurement by recess mounted transducers in a hypersonic boundary layer

Abstract Experimental measurements of pressure fluctuations in a hypersonic boundary layer were conducted based on recess mounted transducers and the effects of recess depth and opening diameter on fluctuating pressure characteristics were studied. The results showed that as the recess depth increased, the bulk of counter flow in the cavity enlarged with the decreased energy of large scale structures, and more small scale structures with relatively high frequency were generated which contributed to the multiscale structures with broadband frequency spectrum. Similarly, as the opening diameter increased, the low-frequency large scale structures inside the cavity had larger energy due to the shear flow entering the cavity having more mass and carrying larger energy, and the small scale structures with high frequency exhibited stronger dissipation intensity. The symmetric and stable flow filed was more likely to cause the cavity resonance, and the asymmetry and instability of the flow field for smaller recess depth as well as larger opening diameter could inhibit the cavity resonance. Compared to surface mounted, nonlinear phase coupling occurred between the flow structures caused by recess mounting which transfered the energy from high-frequency structures to low-frequency structures. On the whole, smaller recess depth and larger opening diameter are optimized in the hypersonic flight test on the premise of pressure sensors not destroyed by the harsh aerothermodynamic environment.

Surface Heating and Boundary-Layer Transition on a Hypersonic Inflatable Aerodynamic Decelerator

The aerothermodynamic environment of a hypersonic inflatable aerodynamic decelerator with a flexible thermal protection system has been investigated through wind-tunnel testing. Boundary-layer transition onset locations and surface heating distributions were measured using global phosphor thermography on deflected, solid-surface models representative of a hypersonic inflatable aerodynamic decelerator aeroshell with a flexible surface. Data were obtained at Mach 6 for a wide range of surface deflection heights, angles of attack, and freestream Reynolds numbers. The surface deflections resulted in earlier boundary-layer transition onset and higher heating levels than for a smooth-wall surface, with the greatest effects occurring on the leeward side of the aeroshell. The heating augmentation and transition onset data were correlated using flowfield properties obtained from a complementary computational study. A method was developed to estimate flexible thermal protection system deflection effects using these correlations along with smooth-wall flowfield predictions. This method was validated through comparisons to the wind-tunnel data and then used to generate postflight predictions for aeroshell heating levels of the Inflatable Reentry Vehicle Experiment flight-test program. The results of this study provide a means for estimating flexible thermal protection system deflection effects for future flight tests as well as an experimental database for use in the development and validation of computational methods for simulations of hypersonic inflatable aerodynamic decelerator aerothermodynamic environments.