Aerothermal Environment Research Articles

Real-Time Aerodynamic Load Estimation for Hypersonics via Strain-Based Inverse Maps

This work develops an efficient real-time inverse formulation for inferring the aerodynamic surface pressures on a hypersonic vehicle from sparse measurements of the structural strain. The approach aims to provide real-time estimates of the aerodynamic loads acting on the vehicle for ground and flight testing, as well as guidance, navigation, and control applications. Specifically, the approach targets hypersonic flight conditions where direct measurement of the surface pressures is challenging due to the harsh aerothermal environment. For problems employing a linear elastic structural model, the inference problem can be posed as a least-squares problem with a linear constraint arising from a finite element discretization of the governing elasticity partial differential equation. Due to the linearity of the problem, an explicit solution is given by the normal equations. Precomputation of the resulting inverse map enables rapid evaluation of the surface pressure and corresponding integrated quantities, such as the force and moment coefficients. The inverse approach additionally allows for uncertainty quantification, providing insights for theoretical recoverability and robustness to sensor noise. Numerical studies demonstrate the estimator performance for reconstructing the surface pressure field, as well as the force and moment coefficients, for the Initial Concept 3.X (IC3X) conceptual hypersonic vehicle.

Multi-physics coupling numerical simulation of variable porosity in reactive melt infiltration process

Abstract The hypersonic craft necessitates significant advancements in thermal shielding to safeguard against the extreme aero-thermal environments encountered during flight at hypersonic speeds in near space or upon reentry from space. Carbon fiber-reinforced ceramic matrix composites (FRCMCs) have attracted increasingly interesting attention as advanced structural materials for application in thermal protection. The reactive melt infiltration (RMI) process is preferred for the preparation of FRCMCs because of its low cost and short processing time. However, the RMI process includes the complex interplay of melt flow, reaction dynamics, and thermal fields, with existing models failing to accurately predict these processes due to their uncoupled nature and neglect of pore structure changes. This study introduces a coupled model that integrates mass transfer flow, chemical reaction, and heat transfer, incorporating variable pore structures to more accurately evaluate the RMI process. This approach allows for a comprehensive assessment of internal pressure, temperature, and porosity changes during silicon infiltration into porous media, marking an improvement over the previous model. The model has been validated for use in predicting the exotherm of RMI processes. In conclusion, this coupled model presents a promising avenue for enhancing the fabrication of FRCMCs, either considering the melt infiltration from multiple locations in the RMI process or adjusting the pore structure at the infiltration location to larger pores.

Cooling film's suppressive effects on infrared system's radiation noise in aero-thermal environment

Cooling film's suppressive effects on infrared system's radiation noise in aero-thermal environment

Pyrolysis front detection in carbon phenolic composites using x-ray computed tomography

Carbon phenolic composites are used as thermal protection systems (TPS) materials on space capsules to protect them from the hot aerothermal environment. The phenolic resin in the composite material decomposes (pyrolyzes) at low temperatures resulting in a pyrolysis front within the material. The detection of the pyrolysis front after exposure to heat has historically been achieved by physically sectioning cross-sections of the material. We combine the phase contrast retrieval method to reconstruct x-ray computed tomography scans along with image convolution to identify the pyrolysis front in carbon phenolic composites. Unlike the standard filtered back projection method that captures only the carbon phase, the phase contrast retrieval method uses both the attenuation coefficients and refractive indices to illuminate all three phases (carbon, resin, and voids) of carbon phenolic composites. Image convolution is applied on scans reconstructed using the phase contrast retrieval method to develop a density map of the composite to locate the pyrolysis front. The analysis is performed on a sample of phenolic impregnated carbon ablator that was tested in an arc-jet facility. For the sample analyzed, the depth of the pyrolysis front from the surface of the sample is calculated to be 2.150 ± 0.148 mm. Although the proposed approach is applied to detect the pyrolysis front, the tools can be used to illuminate the structure of any carbon phenolic composite, and we propose the use of the phase contrast retrieval method as a methodological standard to analyze carbon phenolic composites used on space capsules.

Fabrication and multiscale heat transfer modeling of high thermal conductivity Cf/HfB2-SiC composites

With the increasing Mach numbers of hypersonic vehicles, extreme aerothermal environments impose substantial challenges on vehicle materials. High thermal conductivity materials facilitate rapid heat redistribution to diminish localized thermal accumulation and reduce material thermal damage, providing a viable route for further enhancing the operational temperature of materials. Here, we successfully fabricate high thermal conductivity Cf/HfB2-SiC composites through the solid–liquid combination process, achieving an x-direction thermal conductivity of 161.11 W/(m·K), which represents an order of magnitude improvement over traditional carbon fiber-reinforced ultra-high temperature ceramic composites. 3D X-ray computed tomography (CT) is employed to analyze the structural composition of the high thermal conductivity Cf/HfB2-SiC composites, and the Gaussian Mixture Model is applied to segment the CT images, thus establishing a mesoscale representative volume element (RVE) model based on CT. Combining theoretical models and finite element analysis, we propose a multiscale heat transfer model for high thermal conductivity Cf/HfB2-SiC composites, which effectively investigates the heat transfer mechanisms of the composites. The fabrication and multiscale heat transfer modeling of high thermal conductivity Cf/HfB2-SiC composites provide a theoretical basis and guidance for enhancing material performance and design optimization.

Thermo-mechanical coupling failure mechanisms of reusable SiCf/SiC composites with PyC interphase

The extreme and repeated aerothermal environments working on reusable launch vehicles bring challenges for the design of hot structure. Due to the outstanding high-temperature resistant, SiCf/SiC composite is an excellent candidate material for hot structures in re-entry vehicles. In this paper, the thermo-mechanical coupling failure mechanisms of SiCf/SiC composite are investigated comprehensively. Experimental results reveal different failure mechanisms of SiCf/SiC composites subjected to monotonic quasi-static tension at room temperature and high temperature, cyclic thermal fatigue and thermo-mechanical coupling fatigue. Cyclic thermal environment causing oxidation of interphase layer and embrittlement of fiber plays a significant role in decrease of load bearing capacity of SiCf/SiC composites. The long-term periodic fatigue load that is superimposed on the cyclic thermal condition would further deteriorate the performance of materials. To better understand the damage evolution process, a thermo-mechanical coupling fatigue model is proposed to describe the degradation of SiCf/SiC composites for different stress levels. The service life of SiCf/SiC composites is obtained, which could be used to evaluate the reusable performance of SiCf/SiC hot structure. The results of this study provide novel insights into the design of SiCf/SiC composites for reusable launch vehicles under prolonged and repeated service environment.

A numerical oxidation model for SiC‐coated thermal protection systems in hypersonic applications

AbstractSilicon carbide (SiC) is a widely used component for thermal protection system (TPS) design. This material is often used either as a coating layer to protect beneath carbon‐based material against oxidation or as a ceramic matrix for composite materials used in the hypersonic leading edges. A numerical equilibrium model for SiC‐coated material ablation is developed in this study to compute the material response of coated TPS in hypersonic relevant aerothermal environment. The developed model is based on an equilibrium approach similar to the carbon graphite oxidation framework for ablative charring materials. The equilibrium wall composition is calculated based on the minimization of the Gibbs free energy. The proposed method can capture numerically the transition from passive to active oxidation while maintaining a low computational cost and good numerical stability. Implementation of the model in the Porous Material Analysis Toolbox based on OpenFOAM code permits to compare against previous literature test cases. By comparing the surface temperature obtained from the model with the experimental data, it is observed that the numerical model gives a good estimation providing accurate surface temperature prediction.

Variable Switching System for Heat Protection and Dissipation of Ultra-LEO Satellites Based on LHP Coupled with TEC

Ultra-low Earth orbit (LEO) satellites are widely used in the military, remote sensing, scientific research, and other fields. The ultra-LEO satellite faces the harsh aerothermal environment, and the complex variable attitude task requires the radiator of the satellite to not only meet the heat dissipation requirements of the load but also to resist aerothermal flux. In this study, the aerothermal flux of 160–110 km was calculated, and the loop heat pipe (LHP) coupled with a thermo-electric cooler (TEC) and multi-layer insulation (MLI) were applied to ultra-LEO satellites to determine the variable switching and fast response of heat dissipation and heat protection. An aerothermal flux simulation test platform was built. After the assessment of the ultra-LEO aerothermal flux test, even when the head temperature was as high as 350 °C and the side radiator temperature was as high as 160 °C, the temperature of the internal heat source could be controlled within 22.5 °C through the efficient work of the thermal variable switch system. The study confirms the accuracy and feasibility of the system, which provides an important reference for the subsequent actual on-orbit mission.

Materials design for hypersonics.

Hypersonic vehicles must withstand extreme conditions during flights that exceed five times the speed of sound. These systems have the potential to facilitate rapid access to space, bolster defense capabilities, and create a new paradigm for transcontinental earth-to-earth travel. However, extreme aerothermal environments create significant challenges for vehicle materials and structures. This work addresses the critical need to develop resilient refractory alloys, composites, and ceramics. We will highlight key design principles for critical vehicle areas such as primary structures, thermal protection, and propulsion systems; the role of theory and computation; and strategies for advancing laboratory-scale materials tomanufacturable flight-ready components.

A stability-enhanced approach for aerodynamic heating computation in high-speed flows

Near-space vehicles face severe aerothermal environments due to hypersonic speeds. Accurately predicting hypersonic aerodynamic heating is crucial for the design of spacecraft’s thermal protection systems, aerodynamic shapes, and propulsion systems. Numerical simulations have become an important tool for computing aerothermal environments under extreme conditions. However, there is a contradiction between the robustness and low dissipation characteristics of existing numerical methods for hypersonic flow fields. This paper proposes an effective dimension hybrid approach that enhances the robustness of AUSM-family schemes while maintaining their low dissipation characteristics. Numerical experiments demonstrate the proposed approach’s effectiveness and potential for practical engineering applications.

Stationary Stochastic Response Behaviors of FG Plate–Shell Combined Structure in Aero-Thermal Environment

A meshless model is proposed to carry out the free and stationary stochastic vibration analysis of the functionally graded combined rectangular and cylindrical shells (FG-CRCS) structure under the aerodynamic and thermal environment. The FG-CRCS structure contains three combined structure types including rectangular–cylindrical shell (RC-type), cylindrical–rectangular–cylindrical shell (CRC-type), and closed rectangular–semicylindrical shell (CRS-type). The effects of aerodynamic and thermal loads on the dynamic behaviors of the FG-CRCS structure with temperature-dependency of material properties are investigated by introducing the supersonic piston theory and thermo-elastic theory. Furthermore, the pseudo-excitation method (PEM) is adopted to simulate the random loads applied to the FG-CRCS structure. The dynamic equations of the FG-CRCS structure are established in the theoretical frame of the first-order shear deformation theory (FSDT), whose general boundary conditions and coupling relationship are regulated by the artificial springs. Then, the reasonableness of this meshless model to predict free and random vibrations in aerodynamic and thermal environments is verified by comparing it with published literature and FEM results. On this basis, the contribution of essential parameters (including the aerodynamic load, thermal load, and boundary condition) on the free and random vibration behaviors of the FG-CRCS structure is presented, which may serve as guidance for the design of the plate–shell coupled structures in aerospace.

Mars2020 Entry Shock Layer Thermochemical Kinetics Examined by Megahertz-Rate Laser Absorption Spectroscopy

A mid-infrared laser absorption diagnostic was deployed to examine the evolution of thermophysical properties across a simulated Mars2020 shock layer in the Electric Arc Shock Tube (EAST) facility at NASA Ames. Rapid laser tuning techniques using bias-tee circuitry enabled quantitative temperature and number density measurements of CO2 and CO with microsecond resolution over a shock velocity range of 1.30–3.75 km/s. Two interband cascade lasers were utilized at 4.17 and 4.19 μm to resolve rovibrational CO2 lines spanning across J′′=58 to J′′=140 in the asymmetric stretch fundamental bands. In test cases with enough energy to dissociate CO2, a quantum cascade laser scanned multiple transitions of the CO fundamental bands near 4.98 μm. The results are compared to the Data Parallel Line Relaxation (DPLR) code and Lagrange shock tube analysis (LASTA) simulations of the shock layer. A numerical simulation of the compressible boundary layer is used to account for measurement sensitivities to this flow feature in the EAST facility. Temperature and species transients are compared to multiple chemical kinetic models. The laser absorption data presented in this work can be used to refine the models used to simulate the aerothermal environment encountered during Mars entry.

Realization of electron transpiration cooling: LaB6 heated by a plasma glow

Electron transpiration cooling has previously been predicted to be an alternative cooling mechanism for leading edges of aerospace platforms traveling at extreme velocities, where thermoelectric materials shaped as leading edges manage heat loads. Here, thermo-electron emitting LaB6 ceramics with blunt edges were placed between two conical-shaped electrodes, and a plasma glow was ignited to heat the sample. Analysis of current–voltage (I–V) curves demonstrated that insertion of LaB6 into the plasma column resulted in an increase in the electron density of the plasma, evidencing thermally stimulated electron emission. With increasing sample temperature to 1200 K, Δne increased by ∼4 × 1010 #/cm3, the electron temperature (Te) decreased from ∼8 eV (1000 K) to 2.5 eV, and the emission begins to appear to become space charge-limited. Richardson's analysis of the temperature dependence of the electron emission in the region of ion saturation, yielded an activation energy consistent with the work function of LaB6. Modulation of the LaB6 surface temperature (ΔT = 10 K) and the plasma current (ΔI = 70 μA) under electric voltage (80 V) applied between the sample and current probe was demonstrated. The results show for the first time the feasibility of electron transpiration cooling effects under plasma conditions similar to that of aerothermal environments.

Aftbody Heat Flux Measurements During Mars 2020 Entry

The Mars Entry, Descent, and Landing Instrumentation 2 (MEDLI2) sensor suite on the Mars 2020 mission included two total heat flux sensors and one radiometer on the backshell to directly measure the aftbody aerothermal environments during entry into the Martian atmosphere. All three sensors successfully returned aftbody entry heating measurements. Comparisons between the total heat flux sensor measurements and predictions by NASA simulation tools (DPLR/NEQAIR) show excellent agreement and provide confidence in the models. The radiometer measured significant radiative heating, but compared to the model predictions the signal was attenuated by 48% at the end of the entry heat pulse. The loss of signal is attributed to blockage by thermal protection system (TPS) ablation product deposits on the radiometer window. Ground-based testing in the NASA Ames arcjet facilities was conducted to understand the impact of ablation product deposits on the measured radiometer signal. A discussion of the test results, how flight-like the test conditions were, and future work to further characterize the effect of TPS ablation product deposits on the radiometer performance are presented. In addition to measuring the entry heat pulse, all three sensors were sensitive enough to measure solar radiation during cruise, the radiometer measured solar flux during the entry heat pulse, and the leeside total heat flux sensor picked up the descent reaction control system firings.

Optical performance evaluation of an infrared system of a hypersonic vehicle in an aero-thermal environment.

At hypersonic velocities, the turbulent flow field generated by an aircraft, along with its temperature distribution, leads to significant aerodynamic optical effects that severely impede the performance of internal optical systems. This study proposes a method for analyzing the temporal characteristics of imaging degradation in a detector window infrared imaging system under different field angles of hypersonic velocity. Based on heat transfer theory, a method for solving the transient temperature field in the optical window of a high-speed aircraft is derived and established, considering unsteady thermal conduction-radiation coupling. Additionally, an optical window radiation tracing method is introduced, which directly determines the initial direction vector of light reaching the detector. This method reduces the workload of radiation transmission, significantly enhancing the efficiency of radiation calculations. The time characteristics of image degradation caused by aero-optical effects in high-speed aircraft are analyzed using metrics such as peak signal-to-noise ratio, wave aberration, and point diffusion function. The results demonstrate that as working time increases and the viewing angle widens, the impact of aero-optics on the aircraft imaging system becomes more severe. Moreover, compared to the aerodynamic light transmission effect, the aerodynamic thermal radiation effect has a more detrimental influence on imaging quality.

Heat transfer to proximal cylinders in hypersonic flow

Uncontrolled atmospheric entry of meteors, satellites, and spacecraft components often leads to their partial or complete demise. In this destructive process, driven by hypersonic ablation, reentry objects fragment, interact, and alter each other's aerothermal environment. The effect of these interactions on the heat transfer to the fragments has not been investigated, despite the heat transfer's importance in hypersonic ablation and reentry demise. This study focuses on the numerical investigation of heat transfer to proximal circular cylinders in a thermochemically frozen flow and in two dimensions. First, binary body configurations at Mach numbers 2, 4, and 8 revealed that the heat load and peak heat transfer can be augmented for either or both proximal bodies by +20% to −90% of an isolated body. Second, different clusters of five proximal bodies showed that the heat load to any given body can range from +40% to −90% of an isolated body. Moreover, the average heat load in a cluster is found to vary between +20% and −60% of an isolated body. Intuitively, clusters which are thin in the direction perpendicular to free-stream velocity and long in the direction parallel to the free-stream velocity have their cluster-averaged heat load reduced. In contrast, thick and thin clusters, in directions perpendicular and parallel to the free-stream velocity, are subject to an increased cluster-averaged heat load.

Influence of Ablation Deformation on Aero-Optical Effects in Hypersonic Vehicles

High-speed turbulence is generated when hypersonic vehicles fly in the atmosphere, which can create aero-optical effects and interfere with optical navigation and guidance systems. At the same time, the front end and optical window of hypersonic vehicles are exposed to an aerodynamic heating environment, leading to the head ablation and thermal deformation of the optical window. This further aggravates the turbulent transition process and makes the error of the aero-optical effects more difficult to predict. In this paper, the aero-optical effects under the condition of high-temperature ablation were analyzed. Ablation deformation models of both the head and optical window were established. Then, a high-speed flow field was simulated under different flight conditions. The distortion characteristics of the aero-optical effects were obtained through the photon transmission theory. The simulation results show that the ablation deformation of hypersonic vehicles under an aerodynamic heating environment aggravates the disturbance error of the aero-optical effects. Moreover, with the increase in the flight speed and the decrease in the flight altitude, the ablation deformation of the hypersonic vehicles and the aero-optical effects distortion both gradually increase. The research in this paper provides a reference for the prediction of aero-optical distortion in an aerothermal environment.

Free vibration and random dynamic analyses for the composite cabin-like combined structure in aero-thermal environment

This work presents the numerical investigations on the modal behaviors and stochastic dynamics of a composite cabin-like combined structure considering the aero-thermal factor and various random loads based on the meshless approach. Following the thermo-elastic theory, the influence of the aero-thermal factor on the dynamic behaviors of the combined structure is included. The cabin-like combined structure is disassembled into the open elliptical, conical and cylindrical shells, and coupling conditions between adjacent subshells are imposed by the massless connection springs, in which the Hamilton's principle is adopted to establish free and stochastic vibration models of each subshell. The dynamic solutions for the combined structure system are calculated by employing the meshless technique combined with the pseudo excitation method (PEM). The numerical validity concerning meshless model is presented by performing several case studies related to the modal frequencies and stochastic responses. Additionally, some physical insights into the influences of aero-thermal effect, type of random loads, and geometric dimensions, etc., on the free as well as stochastic vibration properties of the composite cabin-like combined structure are provided.

Influence of Non-Uniform Bluntness on Aerodynamic Performance and Aerothermal Characteristics of Waverider

The waverider is widely used in hypersonic vehicles with its high aerodynamic performance, but due to the serious aerothermal environment, its sharp leading edge should be blunted. Circular blunt is one of the commonly used aerothermal characteristic protection methods. Circular blunt with larger diameter can reduce peak heat flux, but at the same time, it will lead to larger drag. The existing research shows that under the same blunt diameter in two-dimensions, the non-uniform blunt can reduce the peak heat flux by 20%, and the difference of drag is small. In this paper, the non-uniform blunt profile is applied to the three-dimensional waverider, and the influence of the non-uniform blunt profile on the aerothermal characteristic performance and aerodynamic performance of the waverider is studied, and the results are compared with those of circular blunt. The numerical simulation is used to compare and analyze the waverider under different angles of attack, flight altitudes, and Mach number. The results show that the peak heat flux of the waverider with non-uniform blunt reduces by about 17% compared with that with circular blunt under a small angle of attack range, Mach 2-10, and a flight altitude of 15–35 km. Meanwhile, when the blunt height/diameter is 20 mm, the aerodynamic performance difference between the two different blunt profiles does not exceed 3% within a 15 degrees angle of attack, Mach 2-10, and flight altitude of 15–35 km. The non-uniform blunt profile can be applied to the design of the three-dimensional waverider.

Meshless analysis for modal properties and stochastic responses of heated laminated rectangular/sectorial plates in supersonic airflow

The object of this work is to develop a general meshless vibration model for analyzing the modal properties and stochastic responsesof laminated rectangular and sectorial plates subjected to stationary/nonstationary base acceleration random excitation in an aero-thermal environments. The supersonic piston theory is combined with the thermo-elastic theory to consider the thermal and aerodynamic effects of the laminated plate, and the Hamilton's principle is employed for formulating the complete governing equations of the structure system. The displacement components related to dynamic equations and the boundary conditions are characterized by a set of Chebyshev meshless shape functions with orthogonal properties. Then, the vibration behaviors for laminated rectangular/sectorial plates subjected to various boundary constraints are solved, where the pseudo excitation method (PEM) is adopted for performing the stationary/non-stationary stochastic analysis. The accuracy and stability of the present meshless solutions are validated by implementing some convergence checks and verification studies, in which the available results of FEM simulation models as well as the archived literature are provided as reference data. Also, physical insights into the effects of boundary stiffness, thermal/aerodynamic loads and different random excitations on the vibration mechanisms of laminated rectangular/sectorial plates are given, which might serve as the helpful guides for structural dynamic design.