Stagnation Point Heating Research Articles

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.

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.

Stagnation-point heating of Fire II with a non-Boltzmann radiation model

The present work analyzes the stagnation-point radiative heating in the Fire II flight experiment by devising a collisional-radiative model with non-local absorption. In the stagnation-line flow-field calculations, a viscous shock layer method with a thermochemical nonequilibrium model is utilized. In the radiation calculations, a line-by-line method with the non-Boltzmann electronic populations is employed by adopting the quasi-steady state approach of the electronic master equation calculations. In constructing the electronic master equation, the best set of the electron and heavy-particle impact excitation rates is proposed to achieve better agreement with the measured radiative heating flight data. In the flow-radiation coupling procedure, the effect of the non-local absorption is modeled by devising an iterative process between the quasi-steady state electronic master equation and the radiative transfer equation calculations. Escape factors of the strongest atomic lines and the diatomic nitrogen vacuum ultraviolet systems with the non-local absorption effect are also proposed to more efficiently consider the non-local nature of the radiative transition. When compared with the experimental data from the Fire II trajectories, it is found that the present collisional-radiative model with the non-local absorption improves the ability to predict non-Boltzmann radiative heating.

Hypersonic Aeroheating Characteristics of Leading Edges with Different Nose Radii

The applications of leading edges with different nose radii in hypersonic flights require investigations on their aeroheating characteristics in various gas flow regimes. For a detailed numerical study of this problem, the direct simulation Monte Carlo method is used to compute the wall heat fluxes and flowfield information for hypersonic flows over a series of leading-edge profiles. The stagnation point heat flux increases as the nose radius decreases but reaches a finite value when the nose radius approaches zero. This variation characteristic conflicts with the formulas based on continuum flow theory. A further analysis of the flowfield structures illustrates the continuum-breakdown phenomena caused by the strong local nonequilibrium in the vicinity of sharpened leading edges, where a molecular kinetic description of the gas flow is necessary. The distribution functions of molecular velocity at the stagnation points of different leading edges are presented and successfully explain the relation between the stagnation point heating and the nose radius. A fast estimation formula for the stagnation point heat transfer coefficient is proposed. It applies to various flow regimes and gives results that agree well with the present numerical results and the existing experimental data.

Radiative Heating During Mars Science Laboratory Entry: Simulation, Ground Test, and Flight

The heat shield of the Mars Science Laboratory was equipped with thermocouple stacks to measure in-depth heating of the thermal protection system during atmospheric entry. The heat load derived from the thermocouples in the stagnation region was found to be 33% lower than corresponding postflight predictions of convective heating alone. It was hypothesized that this difference could be attributed to radiation from the shock-heated gas, a mechanism not considered in preflight analyses of flowfields. To test the hypothesis and quantify the contribution of shock-layer radiation to total surface heating, ground tests and simulations (both flow and radiation) were performed at several points along the best-estimated entry trajectory of the Mars Science Laboratory. The present paper provides an assessment of the quality of the radiation model and its impact to stagnation point heating. The impact of radiative heating is shown to account for 43% of the heat load discrepancy. Additional possible factors behind the remaining discrepancy are discussed.

Uncertainty Analysis of Mars Entry Flows over a Hypersonic Inflatable Aerodynamic Decelerator

A detailed uncertainty analysis for high-fidelity flowfield simulations over a fixed aeroshell of hypersonic inflatable aerodynamic decelerator scale for Mars entry is presented for fully laminar and turbulent flows at peak stagnation-point heating conditions. This study implements a sparse-collocation approach based on stochastic expansions for efficient and accurate uncertainty quantification under a large number of uncertainty sources in the computational model. The convective and radiative heating and shear stress uncertainties are computed over the hypersonic inflatable aerodynamic decelerator surface and are shown to vary due to a small fraction of 65 flowfield and radiation modeling parameters considered in the uncertainty analysis. The main contributors to the convective heating uncertainty near the stagnation point are the , , and CO–O binary collision interactions, freestream density, and freestream velocity for both boundary-layer flows. In laminar flow, exothermic recombination reactions are more important at the shoulder. The main contributors to radiative heating at the nose and flank were the dissociation rate and CO heavy-particle excitation rates, whereas the freestream density showed importance toward the shoulder. The interaction and freestream velocity and density control the wall shear stress uncertainty.

Variable Transpiration Cooling Effectiveness in Laminar and Turbulent Flows for Hypersonic Vehicles

Reusable thermal-protection systems with active cooling, such as transpiration, are among the promising technologies for thermal management of hypersonic vehicles designed as practical, long-range transportation systems. This paper numerically investigates the effectiveness and efficiency of a variable-velocity transpiration technology for fully laminar and fully turbulent hypersonic flows over a two-dimensional blunt leading-edge geometry. For both flow types, variable transpiration based on a sawtooth velocity distribution is compared to a uniform-velocity transpiration approach. An equal amount of coolant has been imposed to compare the cooling effectiveness between two strategies. The results numerically demonstrate the significant reduction in stagnation-point heating and coolant mass savings achievable with the variable-transpiration strategy, which is observed both in laminar and turbulent flows. The transpiration cooling efficiency is shown to be higher in laminar flow compared to turbulent flow downstream of the leading edge (ramp region). In such regions, for turbulent flows, the amount of total coolant must be increased by a factor of 2 to match the cooling efficiency in laminar flows. The thermal response of a porous thermal-protection-system material is investigated in laminar and turbulent flows with variable transpiration to gain important insight about the matrix and coolant behavior in response to external flow.

Design and Analysis Tool for Mars Atmospheric Entry Missions

A numerical tool is developed for the design and analysis of Mars atmospheric entry trajectories. Simple latitude dependent density model is employed along with an approximate temperature model for Mars' atmosphere. Radiative heat flux and convective heat flux (with cold wall boundary condition) relations are used for the study of stagnation point heating. Approximate drag profile of Mars Pathfinder is used for most simulations. Effects of non-spherical gravity and Mars' surface elevation considerations are discussed in brief. De-orbit thrusting, direct entry, lifting entry, parachute deployment, heat shield release and terminal descent thrusting scenarios are also incorporated into the tool, and their design characteristics are studied. Advantages of lifting entry and the necessity for parachute deployment and terminal descent thrusting are discussed. Different direct entry scenarios are analyzed within feasible launch windows for a typical Mars mission.

Experimental investigations on the effect of energy deposition in hypersonic blunt body flow field

We describe here an experimental study on the effect of energy deposition in the flow field of a $120^0$ blunt cone, carried out in a hypersonic shock tunnel. The energy deposition is realised using an electric arc discharge generated between two electrodes placed in the free stream, and various parameters influencing the effectiveness of this technique is studied. The experimental observations suggest that the location of energy deposition has a vital role in dictating the flow structure, with no noticeable effects being produced on the flow field when the discharge was located close to the body (0.416 times body diameter). In addition, the nature of the test gas and the free stream density are also identified as important parameters. In these experiments, a maximum drag reduction of approxiate \sim50% and \sim84% reduction in stagnation point heating rate has been observed as a result of energy addition. The experimental evidence also indicates that the relaxation of the internal degrees of freedom plays a major role in the alteration of the hypersonic blunt body flow structure and that under the specific conditions encountered in our experiments, the energy deposition is not strong enough to create a shock on its own, but the heated region behind the energy source interacts with the blunt body shock resulting in the flow field alteration.

Gas-surface interaction effect on round leading edge aerothermodynamics

Effects of incomplete surface accommodation in rarefied gas flow have been studied by using the Direct Simulation Monte Carlo (DSMC) method in conjunction with the Cercignani-Lampis-Lord gas surface interaction model. The DSMC calculations examine differences in predictions of aerodynamic forces and heat transfer between full and partial surface accommodation for hypersonic flow past round leading edges at zero incidence. The aerodynamic performance of round leading edges is assessed by using the heat transfer rate, the total drag and the shock wave standoff distance. Twenty-five combinations of nose radius, normal and tangential accommodation coefficients were used in the simulation. For the flow conditions considered, the analysis showed that stagnation point heating, total drag and shock standoff distance are sensitive to changes on either the normal or tangential accommodation coefficient. The results substantiate that it becomes imperative to take surface accommodation into account in order to make accurate predictions of the aerodynamic forces on, and heat transfer rates to, bodies in rarefied hypersonic flow.

Simulation of blunt leading edge aerothermodynamics in rarefied hypersonic flow

The steady-state aerodynamic characteristics of a new family of blunted leading edges immersed in a high-speed rarefied air flow are examined by using a Direct Simulation Monte Carlo (DSMC) Method. A very detailed description of the flow properties has been presented separately at the vicinity of the nose and adjacent to the afterbody surface of the leading edges by a numerical method that properly accounts for non-equilibrium effects that arise near the leading edges and that are especially important at high Mach number. Comparisons based on geometry are made between these new blunt configurations and round leading edge. Some significant differences between these leading edges are noted on the flowfield structure and on the aerodynamic surface quantities. It is found that the upstream effects have different influence on velocity, density, pressure and temperature along the stagnation streamline ahead of the leading edges. The analysis also shows that, despite the seeming advantages of the new blunt leading edge shapes, round leading edge still provides smaller stagnation point heating. Nevertheless, round leading edge provides larger total drag than the new blunt shapes under the range of condition investigated.

Calculation of Stagnation-Point Heating Rates Associated with Stardust Vehicle

Stagnation-point heating rate is calculated for the Stardust entry vehicle and for its heat-shieldmodels tested in an arcjetwind tunnel. The calculations aremade for the convective component for both the laminar and turbulent cases, the radiative component produced by the gas flow, and the radiative component produced by solid particles. A onedimensional viscous shock layer method is used with an up-to-date thermochemical, gas–surface interaction, radiation, turbulence, and particulate models. For the flight cases, radiation absorption in the boundary layer is found to increase the convective heating rate by a factor of up to two. The calculated sums of all heating rates for the flight cases are nearly the same as those obtained by Olynick et al., and are substantially larger than those by Gupta. The calculated sums of all heating rates for the arcjet test cases are smaller than those calculated by the formula of Fay and Riddell, and are nearly the same as those experimentally determined using copper calorimeters.

Arcjet Screening of Candidate Ablative Thermal Protection Materials for Mars Science Laboratory

Arc plasma tests of a range of ablative materials considered candidates for the thermal protection system for the Mars Science Laboratory are described. The objective of these tests was to identify ablative thermal protection materials that offer weight, performance, cost, or other advantages relative to SLA-561V. Several low-density silicones with a mixture of low-density fillers and polymer-impregnated ceramics were tested in the Ames Interaction Heating Facility at conditions representative of nominal stagnation point and worst-case conical heating on the Mars Smart Lander forebody. A wide variety of low-density ablatives and impregnated flexible materials were tested in the Ames Panel Test Facility at nominal and worst-case heating for the Mars Smart Lander backshell. The results of these tests have identified several attractive materials worthy of further investigation.

Parametric Study of Aerocapture for Missions to Venus

A parametric study has been conducted of aerocapture for missions to Venus, examining the entry corridor width as a function of entry velocity, vehicle aerodynamics, and atmospheric density dispersions. The corridor width is found to exceed 1 deg for low lift-to-drag configurations such as an Apollo capsule. For a vehicle with a ballistic coefficient of 78 kg/m 2 and an entry speed of 12 km/s, preliminary calculations indicate that midcorridor trajectories will produce peak stagnation-point heating rates approximately 30% higher than those encountered on Apollo lunar return. Trajectories with modest deceleration loads (under 20 g) and reasonable postaerocapture propulsive requirements were readily identified. Aerocapture is found to merit further study for its potential application on future missions to Venus.

Leading-edge bluntness effects on aerodynamic heating and drag of power law body in low-density hypersonic flow

A numerical study is reported on power law shaped leading edges situated in a rarefied hypersonic flow. The sensitivity of the heat flux and drag coefficient to shape variations of such leading edges is calculated by using a Direct Simulation Monte Carlo method. Calculations show that the stagnation point heating on power law leading edges with finite radius of curvature follows the same relation for classical blunt body in continuum flow; it scales inversely with the square root of the curvature radius at the nose. Furthermore, for those leading edges with zero or infinity radii of curvature, the heat transfer behavior is in surprising agreement with that for classical blunt body far from the nose of the leading edge.

Aerothermodynamic Performance Analysis of Hypersonic Flow on Power Law Leading Edges

A numerical study is reported on power-law-shaped leading edges situated in a rarefied hypersonic flow. The work is motivated by interest in investigating the flowfield properties of power-law-shaped leading edges as possible candidates for blunting geometries of hypersonic leading edges because power-law leading-edge shapes have both blunt (infinite body slope at the nose) and sharp (zero radius of curvature at the nose) characteristics. The sensitivity of the stagnation point heating, total drag, and shock wave standoff distance to shape variations of such leading edges is calculated by using the direct simulation Monte Carlo method. Comparisons based on equivalent stagnation point heating, equivalent total drag, and equivalent shock standoff distance are made between power law shapes and round leading edges. For the flow conditions considered, round leading edges provide lower drag and smaller shock standoff distance than power law leading edges for equivalent stagnation point heating. It was also found that power law shapes have smaller shock standoff distance and larger stagnation point heating than the round leading edges for equivalent total drag.

Wake Closure and Afterbody Heating on a Mars Sample Return Orbiter

Aeroheating wind-tunnel tests were conducted on a 0.028-scale model of an orbiter concept considered for a possible Mars sample return mission. The primary experimental objectives were to characterize hypersonic near-wakeclosure and to determine if shear layer impingement would occur on the proposed orbiter afterbody at incidence angles necessary for a Martian aerocapture maneuver. Global heat transfer mappings, surface streamline patterns, and shock shapes were obtained in the NASA Langley Research Center 20-Inch Mach 6 Air and CF 4 Tunnels for postnormal shock Reynolds numbers (based on forebody diameter) ranging from 1.4 × 10 3 to 4.15 x 105, angles of attack ranging from -5 to 10 deg at 0-, 3-, and 6-deg sideslip, and normal shock density ratios of 5 and 12. Laminar, transitional, and turbulent shear layer impingement on the cylindrical afterbody was inferred from the measurements and resulted in a localized heating maximum that ranged from 40 to 75 % of the reference forebody stagnation point heating.

A simplified trajectory analysis model for small satellite payload recovery from low Earth orbit

The reentry trajectories of a small mass at low Earth orbit were analyzed with a modified three degree-of-freedom trajectory simulation. An estimate of the stagnation point heating was also made. The mass deployed an inflatable, three-element drag disk decelerator that was modeled as a single circular disk, trailing the payload, normal to the direction of flight. The downrange distance decreased and the time of flight increased with increased decelerator area. The stagnation-point heat transfer rates with a decelerator of sufficiently low ballistic coefficient of 24 Pa was 5–10 percent that of a typical ballistic reentry vehicle with a ballistic coefficient of 4.8 kPa. The study did not find that staged deployment of the decelerator disks provided any aerodynamic advantages, particularly in view of the anticipated complexities involved in deploying such a system. Finally, skip trajectories yielded slightly lower stagnation-point heat transfer than the non-lifting case. However, they may not be advantageous since the increased flight time would expose the decelerator to a longer period of heating, thereby possibly requiring more extensive thermal protection and a potential to damage on-board sensors and instruments.

Power-Law Shaped Leading Edges in Rarefied Hypersonic Flow

A numerical study is reported on power-law shaped leading edges situated in a raree ed hypersonic e ow. The work is motivatedby interest in investigating thee owe eld propertiesofpower-lawshapedleading edgesaspossible candidatesforbluntingthegeometryofhypersonicleading edges.Thesensitivityofboth thee owe eld structureand surface quantities to shape variations of such leading edges is calculated by using a direct simulation Monte Carlo method. Particular attention is focused on the pressure gradient along the body surface and on the heat transfer to the body. Calculations showed that the pressure coefe cient for the e ow conditions considered is in surprising agreement with thatobtained by simpleNewtonian theory,which predictsthat a class of power-lawshapes exhibits aerodynamic behavior different from that expected for classical blunt shape. Simulations also showed that the stagnation point heating on power-law leading edges with a e nite radius of curvature follows the same relation as that for classical blunt body. However, for those leading edges with zero radii of curvature, the heat transfer behavior follows another pattern in the vicinity of the stagnation point, reaching a e nite value exactly at the nose.

High enthalpy nozzle flow sensitivity study and effects on heat transfer

The sensitivity of the flow along the nozzle and in the test section of high enthalpy wind tunnels to the thermochemical response of the nozzle expansion process, as well as effects on the pressure and heat transfer distributions over the Electre blunt cone standard test model, are examined in the framework of properly characterizing the test section flow field in such facilities. Particularly sensitive to the thermochemical behaviour of the nozzle flow, in the facilities under consideration, are the static pressure, static temperature and Mach number, whereas stagnation point (pitot) pressure and heat transfer data or freestream velocity are inadequate for the characterization of the thermochemical state of the flow. The Electre and nozzle wall pressure data in the F4 arc jet wind tunnel suggest, in contrast to nonequilibrium computations, that the flow in the F4 nozzle is close to equilibrium. In the HEG and, to some extent, the T5 piston-driven shock tunnels, there are indications that significant heat losses occur in the reservoir. Lastly, simple semi-empirical formulations for stagnation point heating are shown to perform reasonably well in high enthalpy flow conditions.