@inproceedings{
 title = {Assessment of High-Temperature Effects on Hypersonic Aerothermoelastic Analysis using Multi-Fidelity Multi-Variate Surrogates},
 type = {inproceedings},
 year = {2021},
 websites = {https://arc.aiaa.org/doi/10.2514/6.2021-1610},
 month = {1},
 publisher = {AIAA Paper 2021-1610},
 id = {57a31e27-c71f-3461-b901-3f0b3242f497},
 created = {2021-01-05T18:54:29.183Z},
 accessed = {2021-01-05},
 file_attached = {true},
 profile_id = {6476e386-2170-33cc-8f65-4c12ee0052f0},
 last_modified = {2021-04-21T21:17:18.371Z},
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 starred = {false},
 authored = {true},
 confirmed = {false},
 hidden = {false},
 citation_key = {sadagopan:scitech:2021},
 folder_uuids = {5285fb5a-253d-4523-9744-00447d4f0bad,14dd21ec-40bb-405e-90c4-978713737a2f,ad140798-c8ee-4eda-a89e-5d8b660002b6},
 private_publication = {false},
 abstract = {This study investigates the impact of the high-temperature effect, esp. the real gas effect and the chemical reactions, on hypersonic aerothermodynamic solutions of double cone and double wedge configurations, as well as the aerothermoelastic behavior of a double wedge configuration in hypersonic flow. First, a high-temperature computational fluid dynamics code was benchmarked and correlated with experimental results, emphasizing the impact of high-temperature effects as well as turbulence modeling on heat flux prediction. Subsequently, an aerothermal surrogate based on the multi-fidelity Gaussian process regression method was developed. The model achieves a balance between model accuracy and computational cost of sample generation, using the combination of a few high-fidelity sample and many low-fidelity samples. Finally, the new aerothermal surrogate was applied to study the impact of the high-temperature effect on the aerothermoelastic response of a hypersonic skin panel, emphasizing the necessity of the accurate characterization of the localized heat flux for reasonable assessment of the response of a compliant structure in high-speed high-temperature flowfield.},
 bibtype = {inproceedings},
 author = {Sadagopan, Aravinth and Huang, Daning and Martin, Liza E. and Hanquist, Kyle M.},
 doi = {10.2514/6.2021-1610},
 booktitle = {AIAA Scitech 2021 Forum}
}

This study investigates the impact of the high-temperature effect, esp. the real gas effect and the chemical reactions, on hypersonic aerothermodynamic solutions of double cone and double wedge configurations, as well as the aerothermoelastic behavior of a double wedge configuration in hypersonic flow. First, a high-temperature computational fluid dynamics code was benchmarked and correlated with experimental results, emphasizing the impact of high-temperature effects as well as turbulence modeling on heat flux prediction. Subsequently, an aerothermal surrogate based on the multi-fidelity Gaussian process regression method was developed. The model achieves a balance between model accuracy and computational cost of sample generation, using the combination of a few high-fidelity sample and many low-fidelity samples. Finally, the new aerothermal surrogate was applied to study the impact of the high-temperature effect on the aerothermoelastic response of a hypersonic skin panel, emphasizing the necessity of the accurate characterization of the localized heat flux for reasonable assessment of the response of a compliant structure in high-speed high-temperature flowfield.

@inproceedings{
 title = {Effect of Cesium Seeding on Plasma Density in Hypersonic Boundary Layers},
 type = {inproceedings},
 year = {2021},
 websites = {https://arc.aiaa.org/doi/10.2514/6.2021-1251},
 month = {1},
 publisher = {AIAA Paper 2021-1251},
 id = {051c8bad-4a8e-30a6-9913-cafa0a3adcb6},
 created = {2021-01-05T18:58:51.949Z},
 accessed = {2021-01-05},
 file_attached = {true},
 profile_id = {6476e386-2170-33cc-8f65-4c12ee0052f0},
 last_modified = {2021-03-12T00:52:02.831Z},
 read = {false},
 starred = {false},
 authored = {true},
 confirmed = {false},
 hidden = {false},
 citation_key = {parent:scitech:2021},
 folder_uuids = {ce59f875-d211-4f2a-a672-f3278029550e},
 private_publication = {false},
 abstract = {This paper outlines the effect of cesium seeding on the plasma density within the boundary layer around a wedge with a sharp leading edge in the Mach number range 6-18. The results are obtained through numerical simulation using two CFD codes, LeMANS and CFDWARP, which include finite-rate chemistry, non-equilibrium of the vibrational and electron energies, and real gas effects. Results obtained indicate that seeding the air flow with as little as 0.001% of cesium leads to plasma densities high enough to interfere with radio communication and to enable electron transpiration cooling (ETC) at flight Mach numbers as little as 9. When no cesium is added, it is seen that significant interference of the plasma on radio communication can occur in the Mach number range 12-18, with the interference becoming more likely for higher flight dynamic pressure.},
 bibtype = {inproceedings},
 author = {Parent, Bernard and Hanquist, Kyle M. and Rajendran, Prasanna T. and Martin, Liza E.},
 doi = {10.2514/6.2021-1251},
 booktitle = {AIAA Scitech 2021 Forum}
}

This paper outlines the effect of cesium seeding on the plasma density within the boundary layer around a wedge with a sharp leading edge in the Mach number range 6-18. The results are obtained through numerical simulation using two CFD codes, LeMANS and CFDWARP, which include finite-rate chemistry, non-equilibrium of the vibrational and electron energies, and real gas effects. Results obtained indicate that seeding the air flow with as little as 0.001% of cesium leads to plasma densities high enough to interfere with radio communication and to enable electron transpiration cooling (ETC) at flight Mach numbers as little as 9. When no cesium is added, it is seen that significant interference of the plasma on radio communication can occur in the Mach number range 12-18, with the interference becoming more likely for higher flight dynamic pressure.

@inproceedings{
 title = {Effect of cesium seeding on plasma density in hypersonic boundary layers},
 type = {inproceedings},
 year = {2021},
 volume = {1 PartF},
 id = {c961c74d-da41-32e7-984e-47d53c4c32fa},
 created = {2021-02-08T23:59:00.000Z},
 file_attached = {false},
 profile_id = {6476e386-2170-33cc-8f65-4c12ee0052f0},
 last_modified = {2021-04-21T21:17:16.467Z},
 read = {false},
 starred = {false},
 authored = {true},
 confirmed = {false},
 hidden = {false},
 citation_key = {Parent2021},
 folder_uuids = {ad140798-c8ee-4eda-a89e-5d8b660002b6},
 private_publication = {true},
 abstract = {© 2021, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. This paper outlines the effect of cesium seeding on the plasma density within the boundary layer around a wedge with a sharp leading edge in the Mach number range 6–18. The results are obtained through numerical simulation using two CFD codes, LeMANS and CFDWARP, which include finite-rate chemistry, non-equilibrium of the vibrational and electron energies, and real gas effects. Results obtained indicate that seeding the air flow with as little as 0.001% of cesium leads to plasma densities high enough to interfere with radio communication and to enable electron transpiration cooling (ETC) at flight Mach numbers as little as 9. When no cesium is added, it is seen that significant interference of the plasma on radio communication can occur in the Mach number range 12–18, with the interference becoming more likely for higher flight dynamic pressure.},
 bibtype = {inproceedings},
 author = {Parent, B. and Hanquist, K.M. and Rajendran, P.T. and Liza, M.E.},
 booktitle = {AIAA Scitech 2021 Forum}
}

© 2021, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved. This paper outlines the effect of cesium seeding on the plasma density within the boundary layer around a wedge with a sharp leading edge in the Mach number range 6–18. The results are obtained through numerical simulation using two CFD codes, LeMANS and CFDWARP, which include finite-rate chemistry, non-equilibrium of the vibrational and electron energies, and real gas effects. Results obtained indicate that seeding the air flow with as little as 0.001% of cesium leads to plasma densities high enough to interfere with radio communication and to enable electron transpiration cooling (ETC) at flight Mach numbers as little as 9. When no cesium is added, it is seen that significant interference of the plasma on radio communication can occur in the Mach number range 12–18, with the interference becoming more likely for higher flight dynamic pressure.

@inproceedings{
 title = {Modeling High-Temperature Flow Field Effects Relevant to Fluid-Thermal-Structural Interactions},
 type = {inproceedings},
 year = {2020},
 id = {7d567e50-7e32-360c-98f5-8568a4aa830e},
 created = {2020-12-24T00:59:41.671Z},
 file_attached = {false},
 profile_id = {6476e386-2170-33cc-8f65-4c12ee0052f0},
 last_modified = {2021-04-21T21:17:17.403Z},
 read = {false},
 starred = {false},
 authored = {true},
 confirmed = {true},
 hidden = {false},
 citation_key = {hanquist:jannaf:2020},
 source_type = {inproceedings},
 folder_uuids = {5285fb5a-253d-4523-9744-00447d4f0bad,ad140798-c8ee-4eda-a89e-5d8b660002b6},
 private_publication = {false},
 bibtype = {inproceedings},
 author = {Hanquist, Kyle M and Düzel, Ümran and Liza, Martin E and Sadagopan, Aravinth and Huang, Daning},
 booktitle = {Joint Meeting of the Combustion, Airbreathing Propulsion, Exhaust Plume and Signatures, and Energetic Systems Hazards subcommittees, and Programmatic and Industrial Base meeting}
}

@article{
 title = {Shock-tube measurements of coupled vibration-dissociation time-histories and rate parameters in oxygen and argon mixtures from 5000 K to 10 000 K},
 type = {article},
 year = {2020},
 pages = {1-21},
 volume = {32},
 websites = {http://aip.scitation.org/doi/10.1063/5.0012426},
 publisher = {American Institute of Physics Inc.},
 id = {1b5eb4ff-2779-3708-99a5-8b696b23e027},
 created = {2021-01-05T00:36:14.223Z},
 accessed = {2021-01-04},
 file_attached = {true},
 profile_id = {6476e386-2170-33cc-8f65-4c12ee0052f0},
 last_modified = {2021-04-21T21:17:13.896Z},
 read = {false},
 starred = {false},
 authored = {true},
 confirmed = {false},
 hidden = {false},
 citation_key = {streicher:pof:2020},
 folder_uuids = {d365078a-f8a7-46fa-aa09-eac278c395a8,ce59f875-d211-4f2a-a672-f3278029550e,962b6581-dea0-4817-95da-aa57e5178fa0,ad140798-c8ee-4eda-a89e-5d8b660002b6},
 private_publication = {false},
 abstract = {Shock-tube experiments were conducted behind reflected shocks using ultraviolet (UV) laser absorption to measure coupled vibration-dissociation (CVDV) time-histories and rate parameters in dilute mixtures of oxygen (O2) and argon (Ar). Experiments probed 2% and 5% O2 in Ar mixtures for initial post-reflected-shock conditions from 5000 K to 10 000 K and 0.04 atm to 0.45 atm. A tunable, pulsed UV laser absorption diagnostic measured absorbance time-histories from the fourth, fifth, and sixth vibrational levels of the electronic ground state of O2, and experiments were repeated - with closely matched temperature and pressure conditions - to probe absorbance time-histories corresponding to each vibrational level. The absorbance ratio from two vibrational levels, interpreted via an experimentally validated spectroscopic model, determined vibrational temperature time-histories. In contrast, the absorbance involving a single vibrational level determined vibrational-state-specific number density time-histories. These temperature and state-specific number density time-histories agree reasonably well with state-to-state modeling at low temperatures but deviate significantly at high temperatures. Further analysis of the vibrational temperature and number density time-histories isolated coupling parameters from the Marrone and Treanor CVDV model, including vibrational relaxation time (τ), average vibrational energy loss (ϵ), vibrational coupling factor (Z), and dissociation rate constant (kd). The results for τ and kd are consistent with previous results, exhibit low scatter, and - in the case of vibrational relaxation time - extend measurements to higher temperatures than previous experiments. The results for ϵ and Z overlap some common models, exhibit relatively low scatter, and provide novel experimental data.},
 bibtype = {article},
 author = {Streicher, Jesse W. and Krish, Ajay and Hanson, Ronald K. and Hanquist, Kyle M. and Chaudhry, Ross S. and Boyd, Iain D.},
 doi = {10.1063/5.0012426},
 journal = {Physics of Fluids},
 number = {7}
}

Shock-tube experiments were conducted behind reflected shocks using ultraviolet (UV) laser absorption to measure coupled vibration-dissociation (CVDV) time-histories and rate parameters in dilute mixtures of oxygen (O2) and argon (Ar). Experiments probed 2% and 5% O2 in Ar mixtures for initial post-reflected-shock conditions from 5000 K to 10 000 K and 0.04 atm to 0.45 atm. A tunable, pulsed UV laser absorption diagnostic measured absorbance time-histories from the fourth, fifth, and sixth vibrational levels of the electronic ground state of O2, and experiments were repeated - with closely matched temperature and pressure conditions - to probe absorbance time-histories corresponding to each vibrational level. The absorbance ratio from two vibrational levels, interpreted via an experimentally validated spectroscopic model, determined vibrational temperature time-histories. In contrast, the absorbance involving a single vibrational level determined vibrational-state-specific number density time-histories. These temperature and state-specific number density time-histories agree reasonably well with state-to-state modeling at low temperatures but deviate significantly at high temperatures. Further analysis of the vibrational temperature and number density time-histories isolated coupling parameters from the Marrone and Treanor CVDV model, including vibrational relaxation time (τ), average vibrational energy loss (ϵ), vibrational coupling factor (Z), and dissociation rate constant (kd). The results for τ and kd are consistent with previous results, exhibit low scatter, and - in the case of vibrational relaxation time - extend measurements to higher temperatures than previous experiments. The results for ϵ and Z overlap some common models, exhibit relatively low scatter, and provide novel experimental data.

@article{
 title = {Assessment of Thermochemistry Modeling for Hypersonic Flow over a Double Cone},
 type = {article},
 year = {2020},
 pages = {538-547},
 volume = {34},
 websites = {https://arc.aiaa.org/doi/10.2514/1.T5792},
 publisher = {American Institute of Aeronautics and Astronautics Inc.},
 id = {f3ef019c-0853-3b7d-b0cd-835b796d40c0},
 created = {2021-01-05T00:37:42.051Z},
 accessed = {2021-01-04},
 file_attached = {true},
 profile_id = {6476e386-2170-33cc-8f65-4c12ee0052f0},
 last_modified = {2021-04-21T21:17:15.966Z},
 read = {false},
 starred = {false},
 authored = {true},
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 hidden = {false},
 citation_key = {holloway:jtht:2020},
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 private_publication = {false},
 abstract = {The influence of different assumptions for thermochemistry modeling in hypersonic flow over a double-cone geometry is investigated. A computational fluid dynamics analysis is used to study the double cone in three different thermochemical cases, nonequilibrium flow, equilibrium flow, and frozen flow, for four different mixtures of nitrogen and oxygen. Specific areas of interest include the thermochemistry model effects on the flowfield and surface properties. The resulting aerodynamic loads are compared with experiments and indicate that thermochemistry modeling assumptions play a significant role in determining surface properties. It is also shown that heat loading is more sensitive to thermochemical modeling than drag and suggests that an accurate measurement of surface heat transfer is of particular interest. Careful analysis also reveals that high-enthalpy and pure oxygen flows are particularly sensitive to the thermochemistry model assumed. Consistent overprediction or underprediction of pressure drag and heat load by all three chemistry models for most of the cases considered indicates a fundamental difference between the actual experiments and the simulations, thus limiting the usefulness of the double-cone data for validation of thermochemistry models.},
 bibtype = {article},
 author = {Holloway, Michael E. and Hanquist, Kyle M. and Boyd, Iain D.},
 doi = {10.2514/1.T5792},
 journal = {Journal of Thermophysics and Heat Transfer},
 number = {3}
}

The influence of different assumptions for thermochemistry modeling in hypersonic flow over a double-cone geometry is investigated. A computational fluid dynamics analysis is used to study the double cone in three different thermochemical cases, nonequilibrium flow, equilibrium flow, and frozen flow, for four different mixtures of nitrogen and oxygen. Specific areas of interest include the thermochemistry model effects on the flowfield and surface properties. The resulting aerodynamic loads are compared with experiments and indicate that thermochemistry modeling assumptions play a significant role in determining surface properties. It is also shown that heat loading is more sensitive to thermochemical modeling than drag and suggests that an accurate measurement of surface heat transfer is of particular interest. Careful analysis also reveals that high-enthalpy and pure oxygen flows are particularly sensitive to the thermochemistry model assumed. Consistent overprediction or underprediction of pressure drag and heat load by all three chemistry models for most of the cases considered indicates a fundamental difference between the actual experiments and the simulations, thus limiting the usefulness of the double-cone data for validation of thermochemistry models.

@inproceedings{
 title = {Detailed Thermochemical Modeling of O$_2$-Ar in Reflected Shock Tube Flows},
 type = {inproceedings},
 year = {2020},
 publisher = {AIAA Paper 2020-3275},
 id = {6ecbb686-57d8-3aa4-a241-92fc261a0be9},
 created = {2021-01-05T05:00:48.307Z},
 accessed = {2021-01-04},
 file_attached = {true},
 profile_id = {6476e386-2170-33cc-8f65-4c12ee0052f0},
 last_modified = {2021-01-05T05:45:11.266Z},
 read = {false},
 starred = {false},
 authored = {true},
 confirmed = {false},
 hidden = {false},
 citation_key = {hanquist:aviation:2020},
 source_type = {inproceedings},
 private_publication = {false},
 abstract = {Simulation results are presented of a set of vibrational nonequilibrium models with a range of fidelity and are compared to experimental data for several post-normal reflected shock test cases of O2-Ar mixtures. Three different modeling approaches with a range of fidelity are used to determine the vibrational nonequilibrium of the post-normal shock flows. The twotemperature (2T) model is the widely used approach for hypersonic analysis and is presented as the computationally efficient, lower fidelity modeling approach in this work. In contrast, the full state-to-state (STS) model, a master equation approach for each vibrational state, is presented as the higher fidelity modeling approach. Both approaches have several available methods for obtaining rate data that are investigated. The STS approach uses rate data from the forced harmonic oscillator (FHO) approach and quasi-classical trajectory analysis (QCT) for the O2-Ar, O2-O, and O2-O2 systems. The simulated vibrational temperatures and state-specific vibrational level concentrations are compared to experimental measurements. The experimental measurements have a low level of uncertainty and allow for insight into the performance of the nonequilibrium modeling. A rate sensitivity study is also completed that shows how sensitive the results are to certain rates at each experimental condition.},
 bibtype = {inproceedings},
 author = {Hanquist, Kyle M. and Chaudhry, Ross S. and Boyd, Iain D. and Streicher, Jesse W. and Krish, Ajay and Hanson, Ronald K.},
 doi = {10.2514/6.2020-3275},
 booktitle = {AIAA Aviation and Aeronautics Forum and Exposition}
}

Simulation results are presented of a set of vibrational nonequilibrium models with a range of fidelity and are compared to experimental data for several post-normal reflected shock test cases of O2-Ar mixtures. Three different modeling approaches with a range of fidelity are used to determine the vibrational nonequilibrium of the post-normal shock flows. The twotemperature (2T) model is the widely used approach for hypersonic analysis and is presented as the computationally efficient, lower fidelity modeling approach in this work. In contrast, the full state-to-state (STS) model, a master equation approach for each vibrational state, is presented as the higher fidelity modeling approach. Both approaches have several available methods for obtaining rate data that are investigated. The STS approach uses rate data from the forced harmonic oscillator (FHO) approach and quasi-classical trajectory analysis (QCT) for the O2-Ar, O2-O, and O2-O2 systems. The simulated vibrational temperatures and state-specific vibrational level concentrations are compared to experimental measurements. The experimental measurements have a low level of uncertainty and allow for insight into the performance of the nonequilibrium modeling. A rate sensitivity study is also completed that shows how sensitive the results are to certain rates at each experimental condition.

@inproceedings{
 title = {Impact of High-Temperature Effects on the Aerothermoelastic Behavior of Composite Skin Panels in Hypersonic Flow},
 type = {inproceedings},
 year = {2020},
 publisher = {AIAA Paper 2020-0937},
 city = {Orlando, FL},
 id = {19c71fdf-d25f-343b-b52d-7b1d2424eae5},
 created = {2021-01-05T05:00:53.280Z},
 accessed = {2021-01-04},
 file_attached = {true},
 profile_id = {6476e386-2170-33cc-8f65-4c12ee0052f0},
 last_modified = {2021-04-21T21:17:18.838Z},
 read = {false},
 starred = {false},
 authored = {true},
 confirmed = {false},
 hidden = {false},
 citation_key = {sadagopan:scitech:2020},
 source_type = {inproceedings},
 folder_uuids = {5285fb5a-253d-4523-9744-00447d4f0bad,ce59f875-d211-4f2a-a672-f3278029550e,ad140798-c8ee-4eda-a89e-5d8b660002b6},
 private_publication = {false},
 abstract = {This study investigates the impact of the high-temperature effect, esp. the real gas effect and the chemical reactions, on hypersonic aerothermodynamic solutions and the aerothermoelastic behavior of a typical skin panel in hypersonic flow. First, several computational fluid dynamics codes that were developed in significantly different ways were benchmarked and compared for hypersonic aerothermodynamics, emphasizing the impact of high-temperature effects as well as turbulence modeling on heat flux prediction. Subsequently, a reduced-order model (ROM) for hypersonic aerothermal loads accounting for the high-temperature effect is developed. Particularly, a ROM correction approach for high-temperature effect was presented, so that a ROM constructed based on the perfect gas assumption can generate fluid solutions that account for the real gas effect with reasonable accuracy. Finally, the new fluid ROM was applied to study the impact of the high-temperature effect on the aerothermoelastic response of a hypersonic skin panel, with an emphasis on its stability boundary.},
 bibtype = {inproceedings},
 author = {Sadagopan, Aravinth and Huang, Daning and Hanquist, Kyle},
 doi = {10.2514/6.2020-0937},
 booktitle = {AIAA Science and Technology Forum and Exposition}
}

This study investigates the impact of the high-temperature effect, esp. the real gas effect and the chemical reactions, on hypersonic aerothermodynamic solutions and the aerothermoelastic behavior of a typical skin panel in hypersonic flow. First, several computational fluid dynamics codes that were developed in significantly different ways were benchmarked and compared for hypersonic aerothermodynamics, emphasizing the impact of high-temperature effects as well as turbulence modeling on heat flux prediction. Subsequently, a reduced-order model (ROM) for hypersonic aerothermal loads accounting for the high-temperature effect is developed. Particularly, a ROM correction approach for high-temperature effect was presented, so that a ROM constructed based on the perfect gas assumption can generate fluid solutions that account for the real gas effect with reasonable accuracy. Finally, the new fluid ROM was applied to study the impact of the high-temperature effect on the aerothermoelastic response of a hypersonic skin panel, with an emphasis on its stability boundary.

@inproceedings{
 title = {Fully-Coupled Simulation of Plasma Discharges, Turbulence, and Combustion in a Scramjet Combustor},
 type = {inproceedings},
 year = {2020},
 publisher = {AIAA Paper 2020-3230},
 id = {c22247a3-51e1-373a-bc92-e8f1e4f5d1b0},
 created = {2021-01-05T05:02:09.628Z},
 accessed = {2021-01-04},
 file_attached = {true},
 profile_id = {6476e386-2170-33cc-8f65-4c12ee0052f0},
 last_modified = {2021-04-21T21:17:17.867Z},
 read = {false},
 starred = {false},
 authored = {true},
 confirmed = {false},
 hidden = {false},
 citation_key = {parent:avi:2020},
 source_type = {inproceedings},
 folder_uuids = {ad140798-c8ee-4eda-a89e-5d8b660002b6},
 private_publication = {false},
 abstract = {Simulating plasma-assisted combustion represents a considerable challenge due to the large discrepancy of the time scales involved. While the turbulent eddy time scales are of the order of microseconds, the plasma sheath time scales are 3-4 orders of magnitude lower. Contrarily to the chemical reactions, the stiffness of the plasma equations can not be relieved simply by using an implicit integration strategy, thus leading to excessive computational effort even for the simplest cases. Recently, it was shown that this hurdle can be overcome by recasting the plasma driftdiffusion transport equations such that the potential is not obtained from Gauss’s law directly but rather from Ohm’s law. Such a recast is performed while still ensuring that Gauss’s law is satisfied and thus does not modify the physics of the drift-diffusion model in any way. In this paper, we use this novel approach to integrate, for the first time, a plasma discharge in fully coupled form with the turbulent hydrogen/air mixing layer and combustion process taking place in the combustor of a scramjet flying at Mach 11. The chemical model includes electrons, 7 different types of ions, 11 neutral species and 79 reactions. Results indicate that more than 5 discharges need to be performed before achieving a self-repeating pattern due to the strong coupling between the flow, combustion, and plasma. Further, the plasma-assisted flame anchoring is seen to create a recirculation region of significant size within the turbulent boundary layer which affects skin friction and heat loads considerably.},
 bibtype = {inproceedings},
 author = {Parent, Bernard and Omprakas, Ajjay and Hanquist, Kyle M.},
 doi = {10.2514/6.2020-3230},
 booktitle = {AIAA Aviation and Aeronautics Forum and Exposition}
}

Simulating plasma-assisted combustion represents a considerable challenge due to the large discrepancy of the time scales involved. While the turbulent eddy time scales are of the order of microseconds, the plasma sheath time scales are 3-4 orders of magnitude lower. Contrarily to the chemical reactions, the stiffness of the plasma equations can not be relieved simply by using an implicit integration strategy, thus leading to excessive computational effort even for the simplest cases. Recently, it was shown that this hurdle can be overcome by recasting the plasma driftdiffusion transport equations such that the potential is not obtained from Gauss’s law directly but rather from Ohm’s law. Such a recast is performed while still ensuring that Gauss’s law is satisfied and thus does not modify the physics of the drift-diffusion model in any way. In this paper, we use this novel approach to integrate, for the first time, a plasma discharge in fully coupled form with the turbulent hydrogen/air mixing layer and combustion process taking place in the combustor of a scramjet flying at Mach 11. The chemical model includes electrons, 7 different types of ions, 11 neutral species and 79 reactions. Results indicate that more than 5 discharges need to be performed before achieving a self-repeating pattern due to the strong coupling between the flow, combustion, and plasma. Further, the plasma-assisted flame anchoring is seen to create a recirculation region of significant size within the turbulent boundary layer which affects skin friction and heat loads considerably.

@article{
 title = {Aerothermodynamic Design Optimization of Hypersonic Vehicles},
 type = {article},
 year = {2019},
 pages = {392-406},
 volume = {33},
 websites = {https://arc.aiaa.org/doi/10.2514/1.T5523},
 publisher = {American Institute of Aeronautics and Astronautics Inc.},
 id = {ac8fb95b-e361-3caf-85df-d713f8e45568},
 created = {2021-01-05T00:32:19.261Z},
 accessed = {2021-01-04},
 file_attached = {true},
 profile_id = {6476e386-2170-33cc-8f65-4c12ee0052f0},
 last_modified = {2021-04-21T21:17:15.013Z},
 read = {false},
 starred = {false},
 authored = {true},
 confirmed = {false},
 hidden = {false},
 citation_key = {eyi:jtht:2019},
 folder_uuids = {d365078a-f8a7-46fa-aa09-eac278c395a8,5285fb5a-253d-4523-9744-00447d4f0bad,14dd21ec-40bb-405e-90c4-978713737a2f,ad140798-c8ee-4eda-a89e-5d8b660002b6},
 private_publication = {false},
 abstract = {The objective of this study is to develop a reliable and efficient design optimization method for hypersonic vehicles focused on aerothermodynamic environments. Considering the nature of hypersonic flight, a high-fidelity aerothermodynamic analysis code is used for the simulation of weakly ionized hypersonic flows in thermochemical nonequilibrium. A gradient-based method is implemented for optimization. Bezier or nonuniform rational basis spline curves are used to parametrize the geometry or the geometry change. Linear elasticity theory is implemented for mesh deformation. Penalty functions are utilized to prevent undesired geometrical changes. The design objective is to minimize drag without increasing the total heat transfer rate and the maximum values of the surface heat flux, temperature, and pressure. Design optimizations are performed at different trajectory points of the IRV-2 vehicle. The effects of parametrizations, the number of design variables, and freestream conditions on design performance are studied.},
 bibtype = {article},
 author = {Eyi, Sinan and Hanquist, Kyle M. and Boyd, Iain D.},
 doi = {10.2514/1.T5523},
 journal = {Journal of Thermophysics and Heat Transfer},
 number = {2}
}

The objective of this study is to develop a reliable and efficient design optimization method for hypersonic vehicles focused on aerothermodynamic environments. Considering the nature of hypersonic flight, a high-fidelity aerothermodynamic analysis code is used for the simulation of weakly ionized hypersonic flows in thermochemical nonequilibrium. A gradient-based method is implemented for optimization. Bezier or nonuniform rational basis spline curves are used to parametrize the geometry or the geometry change. Linear elasticity theory is implemented for mesh deformation. Penalty functions are utilized to prevent undesired geometrical changes. The design objective is to minimize drag without increasing the total heat transfer rate and the maximum values of the surface heat flux, temperature, and pressure. Design optimizations are performed at different trajectory points of the IRV-2 vehicle. The effects of parametrizations, the number of design variables, and freestream conditions on design performance are studied.

@article{
 title = {Shape Optimization of Reentry Vehicles to Minimize Heat Loading},
 type = {article},
 year = {2019},
 pages = {785-796},
 volume = {33},
 websites = {https://arc.aiaa.org/doi/10.2514/1.T5705},
 publisher = {American Institute of Aeronautics and Astronautics Inc.},
 id = {055faa2d-738a-3a9a-99ad-71858c1fc261},
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 private_publication = {false},
 abstract = {The objective of the current study is to designanoptimumreentry vehicle shape thatminimizes heat loading subject to constraints on themaximumvalues of surface heat flux and temperature. A new heat loading formulation is developed for objective function evaluations. Axisymmetric Navier-Stokes and finite-rate chemical reaction equations are solved to evaluate the objectiveandconstraint functions.TheMenterSSTturbulencemodel isemployedfor turbulence closure. A gradient-based method is used for optimization. The sensitivities of the objective and constraint functions are evaluated using the finite-difference method. In shape optimization, the geometry change or the geometry itself is parameterized using different numbers of nonuniform rational basis spline (NURBS) or Bezier curves. Designs are performed at different trajectory points of the IRV-2 vehicle. The effects of flight path angle and reentry velocity on the heat transfer and trajectory characteristics of the original and designed geometries are quantified.},
 bibtype = {article},
 author = {Eyi, Sinan and Hanquist, Kyle M. and Boyd, Iain D.},
 doi = {10.2514/1.T5705},
 journal = {Journal of Thermophysics and Heat Transfer},
 number = {3}
}

The objective of the current study is to designanoptimumreentry vehicle shape thatminimizes heat loading subject to constraints on themaximumvalues of surface heat flux and temperature. A new heat loading formulation is developed for objective function evaluations. Axisymmetric Navier-Stokes and finite-rate chemical reaction equations are solved to evaluate the objectiveandconstraint functions.TheMenterSSTturbulencemodel isemployedfor turbulence closure. A gradient-based method is used for optimization. The sensitivities of the objective and constraint functions are evaluated using the finite-difference method. In shape optimization, the geometry change or the geometry itself is parameterized using different numbers of nonuniform rational basis spline (NURBS) or Bezier curves. Designs are performed at different trajectory points of the IRV-2 vehicle. The effects of flight path angle and reentry velocity on the heat transfer and trajectory characteristics of the original and designed geometries are quantified.

@article{
 title = {Plasma Assisted Cooling of Hot Surfaces on Hypersonic Vehicles},
 type = {article},
 year = {2019},
 keywords = {etc},
 pages = {1-13},
 volume = {7},
 websites = {https://www.frontiersin.org/article/10.3389/fphy.2019.00009/full},
 publisher = {Frontiers},
 id = {460cc4e0-1a8b-3c2b-9508-b01a896191c8},
 created = {2021-01-05T00:34:23.113Z},
 accessed = {2021-01-04},
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 profile_id = {6476e386-2170-33cc-8f65-4c12ee0052f0},
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 private_publication = {false},
 abstract = {Electron transpiration cooling (ETC) is a proposed thermal management approach for the leading edges of hypersonic vehicles that utilizes thermionic emission to emit electrons to carry heat away from the surface. A modeling approach is presented for assessing ETC in a computational fluid dynamics (CFD) framework and is evaluated using previously completed experiments. The modeling approach presented includes developing boundary conditions to account for space-charge-limited emission to accurately determine the level of electron emission from the surface. The effectiveness of ETC for multiple test cases are investigated including sharp leading edges and blunt bodies. For each of these test cases, ETC affects the surface properties, most notably the surface temperature, suggesting that ETC occurs for bodies in thermally intense, ionized flows, no matter the shape of the leading edge. An approximate approach is also presented to assess ETC in an ionized flow and compares its cooling power to radiative cooling.},
 bibtype = {article},
 author = {Hanquist, Kyle M. and Boyd, Iain D.},
 doi = {10.3389/fphy.2019.00009},
 journal = {Frontiers in Physics: Plasma for Aerospace},
 number = {9}
}

Electron transpiration cooling (ETC) is a proposed thermal management approach for the leading edges of hypersonic vehicles that utilizes thermionic emission to emit electrons to carry heat away from the surface. A modeling approach is presented for assessing ETC in a computational fluid dynamics (CFD) framework and is evaluated using previously completed experiments. The modeling approach presented includes developing boundary conditions to account for space-charge-limited emission to accurately determine the level of electron emission from the surface. The effectiveness of ETC for multiple test cases are investigated including sharp leading edges and blunt bodies. For each of these test cases, ETC affects the surface properties, most notably the surface temperature, suggesting that ETC occurs for bodies in thermally intense, ionized flows, no matter the shape of the leading edge. An approximate approach is also presented to assess ETC in an ionized flow and compares its cooling power to radiative cooling.

@inproceedings{
 title = {Modeling of Electronically Excited Oxygen in O$_2$-Ar Shock Tube Studies},
 type = {inproceedings},
 year = {2019},
 month = {6},
 publisher = {AIAA Paper 2019-3567},
 day = {17},
 city = {Atlanta, GA},
 id = {9ce9ffaf-c797-33f8-a440-83c0ffcdec19},
 created = {2021-01-05T04:58:05.153Z},
 accessed = {2021-01-04},
 file_attached = {true},
 profile_id = {6476e386-2170-33cc-8f65-4c12ee0052f0},
 last_modified = {2021-01-05T05:45:11.175Z},
 read = {false},
 starred = {false},
 authored = {true},
 confirmed = {false},
 hidden = {false},
 citation_key = {hanquist:aviation:2019},
 source_type = {inproceedings},
 private_publication = {false},
 abstract = {The successful development of hypersonic vehicles requires detailed knowledge of the flow physics around the vehicle. Specifically, an understanding of the thermochemical nonequi-librium behavior is crucial for this flight regime. Additionally, the hypersonic flight regime involves an extremely high level of energy so a small error in the modeling of the energy processes can result in drastic changes in the vehicle design, which motivates modeling the physics involved at a high-fidelity. However, there is limited experimental data to assess the current nonequilibrium modeling approaches. Recently, the Hanson Group at Stanford University measured the formation of electronically excited atomic oxygen behind reflected shock waves using cavity-enhanced absorption spectroscopy. The motivation of this work is to develop a modeling approach that can be assessed using these experiments. In the present work, 1D post normal shock flow calculations of both pure and diluted molecular oxygen in argon are carried out and used to analyze existing shock tube experiments. State-of-the-art thermochemical nonequi-librium models, including two-temperature (2T) and multitemperature-collisional-radiative (MTCR) models are adopted in these post normal shock flow analyses. The 2T approach models the excited states using Boltzmann statistics at the vibrational temperature. The MTCR uses a four temperature approach (translational, rotational, vibrational, and electronic). The non-Boltzmann behavior of the excited states is modeled by including the relevant collisional and radiative mechanisms and then solving for the excited state concentrations using an electronic master equation coupling model. I. Nomenclature c = Thermal velocity D = Dissociation energy e = Specific energy E = Energy g = Degeneracy h = Enthalpy I = Ionization energy k = Boltzmann's constant K = Excitation rate m = Mass N a = Avogadro's number n = Number density p = Pressure Q = Partition function Q r ad = Radiative energy loss q = Franck-Condon factor t = Time T = Temperature u = Velocity v = Collision frequency x = Distance from shock * Research Fellow and Lecturer, Member AIAA.},
 bibtype = {inproceedings},
 author = {Hanquist, Kyle M. and Boyd, Iain D.},
 doi = {10.2514/6.2019-3567},
 booktitle = {AIAA Aviation and Aeronautics Forum and Exposition}
}

The successful development of hypersonic vehicles requires detailed knowledge of the flow physics around the vehicle. Specifically, an understanding of the thermochemical nonequi-librium behavior is crucial for this flight regime. Additionally, the hypersonic flight regime involves an extremely high level of energy so a small error in the modeling of the energy processes can result in drastic changes in the vehicle design, which motivates modeling the physics involved at a high-fidelity. However, there is limited experimental data to assess the current nonequilibrium modeling approaches. Recently, the Hanson Group at Stanford University measured the formation of electronically excited atomic oxygen behind reflected shock waves using cavity-enhanced absorption spectroscopy. The motivation of this work is to develop a modeling approach that can be assessed using these experiments. In the present work, 1D post normal shock flow calculations of both pure and diluted molecular oxygen in argon are carried out and used to analyze existing shock tube experiments. State-of-the-art thermochemical nonequi-librium models, including two-temperature (2T) and multitemperature-collisional-radiative (MTCR) models are adopted in these post normal shock flow analyses. The 2T approach models the excited states using Boltzmann statistics at the vibrational temperature. The MTCR uses a four temperature approach (translational, rotational, vibrational, and electronic). The non-Boltzmann behavior of the excited states is modeled by including the relevant collisional and radiative mechanisms and then solving for the excited state concentrations using an electronic master equation coupling model. I. Nomenclature c = Thermal velocity D = Dissociation energy e = Specific energy E = Energy g = Degeneracy h = Enthalpy I = Ionization energy k = Boltzmann's constant K = Excitation rate m = Mass N a = Avogadro's number n = Number density p = Pressure Q = Partition function Q r ad = Radiative energy loss q = Franck-Condon factor t = Time T = Temperature u = Velocity v = Collision frequency x = Distance from shock * Research Fellow and Lecturer, Member AIAA.

@inproceedings{
 title = {Effect of Thermochemistry Modeling on Hypersonic Flow Over a Double Cone},
 type = {inproceedings},
 year = {2019},
 publisher = {AIAA Paper 2019-2281},
 city = {San Diego, CA},
 id = {34691465-7603-37af-a777-a83d59543bfc},
 created = {2021-01-05T04:58:16.450Z},
 accessed = {2021-01-04},
 file_attached = {true},
 profile_id = {6476e386-2170-33cc-8f65-4c12ee0052f0},
 last_modified = {2021-01-05T05:45:11.928Z},
 read = {false},
 starred = {false},
 authored = {true},
 confirmed = {false},
 hidden = {false},
 citation_key = {holloway:scitech:2019},
 source_type = {inproceedings},
 private_publication = {false},
 abstract = {The influence of different assumptions for thermochemistry modeling in hypersonic flow over a double-cone geometry is investigated. The double-cone geometry is simple but produces a complex shock wave/boundary layer interaction and nonequilibrium flow physics. This interaction sig-nificantly impacts the aerothermodynamic loading, in terms of surface pressure and heat transfer. Therefore, it is important that these interactions can be predicted with physical accuracy and numerical efficiency. A CFD analysis is used to study the double-cone in three different thermochemical cases: nonequilibrium flow, equilibrium flow, and frozen flow for five different mixtures of nitrogen and oxygen. Specific areas of interest include the thermochemistry model effects on the flow field and surface properties. The resulting aerodynamic loads are compared to experiments and indicate that thermochemistry modeling assumptions play a significant role in determining surface properties. It is also shown that heat loading is more sensitive to thermochemical modeling than drag and suggests that an accurate measurement of surface heat transfer is of particular interest. Careful analysis also reveals that high enthalpy and pure oxygen flows are particularly sensitive to the thermochemistry model assumed.},
 bibtype = {inproceedings},
 author = {Holloway, Michael E. and Hanquist, Kyle M. and Boyd, Iain D.},
 doi = {10.2514/6.2019-2281},
 booktitle = {AIAA Science and Technology Forum and Exposition}
}

The influence of different assumptions for thermochemistry modeling in hypersonic flow over a double-cone geometry is investigated. The double-cone geometry is simple but produces a complex shock wave/boundary layer interaction and nonequilibrium flow physics. This interaction sig-nificantly impacts the aerothermodynamic loading, in terms of surface pressure and heat transfer. Therefore, it is important that these interactions can be predicted with physical accuracy and numerical efficiency. A CFD analysis is used to study the double-cone in three different thermochemical cases: nonequilibrium flow, equilibrium flow, and frozen flow for five different mixtures of nitrogen and oxygen. Specific areas of interest include the thermochemistry model effects on the flow field and surface properties. The resulting aerodynamic loads are compared to experiments and indicate that thermochemistry modeling assumptions play a significant role in determining surface properties. It is also shown that heat loading is more sensitive to thermochemical modeling than drag and suggests that an accurate measurement of surface heat transfer is of particular interest. Careful analysis also reveals that high enthalpy and pure oxygen flows are particularly sensitive to the thermochemistry model assumed.

@inproceedings{
 title = {Shape Optimization of Reentry Vehicles to Minimize Heat Loading},
 type = {inproceedings},
 year = {2019},
 publisher = {AIAA Paper 2019-0973},
 city = {San Diego, CA},
 id = {791a0138-4886-3635-b2c2-a74fb4ae6b3b},
 created = {2021-01-05T04:59:06.355Z},
 accessed = {2021-01-04},
 file_attached = {true},
 profile_id = {6476e386-2170-33cc-8f65-4c12ee0052f0},
 last_modified = {2021-01-05T05:45:11.525Z},
 read = {false},
 starred = {false},
 authored = {true},
 confirmed = {false},
 hidden = {false},
 citation_key = {eyi:scitech:2019},
 source_type = {inproceedings},
 private_publication = {false},
 abstract = {The objective of the current study is to design an optimum reentry vehicle shape that minimizes heat loading subject to constraints on the maximum values of surface heat flux and temperature. A new formulation is developed for the heat loading calculations. Axisymmetric Navier-Stokes and finite rate chemical reaction equations are solved to evaluate the objective and constraint functions. The Menter SST turbulence model is employed for turbulence closure. A gradient-based method is utilized for optimization. The sensitivities of the objective and constraint functions are evaluated using the finite difference method. In design optimization, the geometry change or the geometry itself is parameterized using different numbers of NURBS or Bezier curves. Designs are performed at different trajectory points of the IRV-2 vehicle. The effects of flight path angle and reentry velocity on the heat transfer and trajectory characteristics of the original and designed geometries are quantified.},
 bibtype = {inproceedings},
 author = {Eyi, Sinan and Hanquist, Kyle M. and Boyd, Iain D.},
 doi = {10.2514/6.2019-0973},
 booktitle = {AIAA Science and Technology Forum and Exposition}
}

The objective of the current study is to design an optimum reentry vehicle shape that minimizes heat loading subject to constraints on the maximum values of surface heat flux and temperature. A new formulation is developed for the heat loading calculations. Axisymmetric Navier-Stokes and finite rate chemical reaction equations are solved to evaluate the objective and constraint functions. The Menter SST turbulence model is employed for turbulence closure. A gradient-based method is utilized for optimization. The sensitivities of the objective and constraint functions are evaluated using the finite difference method. In design optimization, the geometry change or the geometry itself is parameterized using different numbers of NURBS or Bezier curves. Designs are performed at different trajectory points of the IRV-2 vehicle. The effects of flight path angle and reentry velocity on the heat transfer and trajectory characteristics of the original and designed geometries are quantified.

@article{
 title = {Test cases for grid-based direct kinetic modeling of plasma flows},
 type = {article},
 year = {2018},
 keywords = {etc},
 pages = {65004},
 volume = {27},
 websites = {https://doi.org/10.1088/1361-6595/aac6b9},
 publisher = {Institute of Physics Publishing},
 id = {a9a68a9e-2f34-32f0-a56c-1121dc624cc5},
 created = {2021-01-05T00:18:01.001Z},
 accessed = {2021-01-04},
 file_attached = {true},
 profile_id = {6476e386-2170-33cc-8f65-4c12ee0052f0},
 last_modified = {2021-04-21T21:17:14.379Z},
 read = {false},
 starred = {false},
 authored = {true},
 confirmed = {false},
 hidden = {false},
 citation_key = {hara:psst:2018},
 folder_uuids = {d365078a-f8a7-46fa-aa09-eac278c395a8,ce59f875-d211-4f2a-a672-f3278029550e,962b6581-dea0-4817-95da-aa57e5178fa0,ad140798-c8ee-4eda-a89e-5d8b660002b6},
 private_publication = {false},
 abstract = {Grid-based kinetic models are promising in that the numerical noise inherent in particle-based methods is essentially eliminated. Here, we call such grid-based techniques a direct kinetic (DK) model. Velocity distribution functions are directly obtained by solving kinetic equations, such as the Vlasov equation, in discretized phase space, i.e., both physical and velocity space. In solving the kinetic equations that are hyperbolic partial differential equations, we employ a conservative, positivity-preserving numerical scheme, which is necessary for robust calculations of problems particularly including ionization. Test cases described in this paper include plasma sheaths with electron emission and injection and expansion of neutral atom flow in a two-dimensional configuration. A unifying kinetic theory of space charge limited sheaths for both floating and conducting surfaces is presented. The improved theory is verified using the collisionless DK simulation, particularly for small sheath potentials that particle-based kinetic simulations may struggle due to statistical noise. For benchmarking of the grid-based and particle-based kinetic simulations, hybrid simulations of Hall thruster discharge plasma are performed. While numerical diffusion occurs in the phase space in the DK simulation, ionization oscillations are well resolved since ionization events can be taken into account deterministically at every time step.},
 bibtype = {article},
 author = {Hara, Kentaro and Hanquist, Kyle M.},
 doi = {10.1088/1361-6595/aac6b9},
 journal = {Plasma Sources Science and Technology},
 number = {6}
}

Grid-based kinetic models are promising in that the numerical noise inherent in particle-based methods is essentially eliminated. Here, we call such grid-based techniques a direct kinetic (DK) model. Velocity distribution functions are directly obtained by solving kinetic equations, such as the Vlasov equation, in discretized phase space, i.e., both physical and velocity space. In solving the kinetic equations that are hyperbolic partial differential equations, we employ a conservative, positivity-preserving numerical scheme, which is necessary for robust calculations of problems particularly including ionization. Test cases described in this paper include plasma sheaths with electron emission and injection and expansion of neutral atom flow in a two-dimensional configuration. A unifying kinetic theory of space charge limited sheaths for both floating and conducting surfaces is presented. The improved theory is verified using the collisionless DK simulation, particularly for small sheath potentials that particle-based kinetic simulations may struggle due to statistical noise. For benchmarking of the grid-based and particle-based kinetic simulations, hybrid simulations of Hall thruster discharge plasma are performed. While numerical diffusion occurs in the phase space in the DK simulation, ionization oscillations are well resolved since ionization events can be taken into account deterministically at every time step.

@inproceedings{
 title = {Modeling of Excited Oxygen in Post Normal Shock Waves},
 type = {inproceedings},
 year = {2018},
 publisher = {AIAA Paper 2018-3769},
 city = {Atlanta, GA},
 id = {43c88f53-28b7-3291-ac3d-da39c9fce0ae},
 created = {2021-01-05T04:55:00.533Z},
 accessed = {2021-01-04},
 file_attached = {true},
 profile_id = {6476e386-2170-33cc-8f65-4c12ee0052f0},
 last_modified = {2021-01-05T05:45:11.114Z},
 read = {false},
 starred = {false},
 authored = {true},
 confirmed = {false},
 hidden = {false},
 citation_key = {hanquist:aviation:18},
 source_type = {inproceedings},
 private_publication = {false},
 abstract = {The successful development of hypersonic vehicles requires a detailed knowledge of the flow physics around the vehicle. Specifically, an understanding of the thermochemical nonequilibrium behavior is crucial for this flight regime. The hypersonic flight regime involves an extremely high level of energy, so a small error in the modeling of the energy processes can result in drastic changes in the vehicle design, which motivates modeling the physics involved at a high-fidelity. Recent progress is presented in an ongoing effort to model the excited states of oxygen in post-normal shock waves using computational fluid dynamics. One-dimensional post normal shock flow calculations are carried out using state-of-the-art thermochemical nonequilibrium models. Two-temperature and electronic master equation coupling models are adopted in the present work and discussed in detail. Different approaches of modeling the energy transfer from each mode are also presented. The approaches are assessed using a set of existing experiments where the vibrational temperature was measured. The concentrations of excited states of atomic oxygen determined by the electronic master equation coupling model are compared to Boltzmann distributions.},
 bibtype = {inproceedings},
 author = {Hanquist, Kyle M. and Boyd, Iain D.},
 doi = {10.2514/6.2018-3769},
 booktitle = {2018 Joint Thermophysics and Heat Transfer Conference}
}

The successful development of hypersonic vehicles requires a detailed knowledge of the flow physics around the vehicle. Specifically, an understanding of the thermochemical nonequilibrium behavior is crucial for this flight regime. The hypersonic flight regime involves an extremely high level of energy, so a small error in the modeling of the energy processes can result in drastic changes in the vehicle design, which motivates modeling the physics involved at a high-fidelity. Recent progress is presented in an ongoing effort to model the excited states of oxygen in post-normal shock waves using computational fluid dynamics. One-dimensional post normal shock flow calculations are carried out using state-of-the-art thermochemical nonequilibrium models. Two-temperature and electronic master equation coupling models are adopted in the present work and discussed in detail. Different approaches of modeling the energy transfer from each mode are also presented. The approaches are assessed using a set of existing experiments where the vibrational temperature was measured. The concentrations of excited states of atomic oxygen determined by the electronic master equation coupling model are compared to Boltzmann distributions.

@inproceedings{
 title = {Effectiveness of Thermionic Emission for Cooling Hypersonic Vehicle Surfaces},
 type = {inproceedings},
 year = {2018},
 publisher = {AIAA Paper 2018-1714},
 city = {Kissimmee, F},
 id = {6e2cc7fa-4e1e-393b-b0a0-ac3c2869e40a},
 created = {2021-01-05T04:55:39.295Z},
 accessed = {2021-01-04},
 file_attached = {true},
 profile_id = {6476e386-2170-33cc-8f65-4c12ee0052f0},
 last_modified = {2021-01-05T05:45:11.494Z},
 read = {false},
 starred = {false},
 authored = {true},
 confirmed = {false},
 hidden = {false},
 citation_key = {hanquist:scitech:2018},
 source_type = {inproceedings},
 private_publication = {false},
 abstract = {Electron transpiration cooling (ETC) is a proposed thermal management approach for the leading edges of hypersonic vehicles that utilizes thermionic emission to emit electrons to carry heat away from the surface. This paper presents a modeling approach for implementing ETC in a computational fluid dynamics (CFD) framework and assesses the modeling approach using a set of previously completed experiments. The modeling approach includes coupling the fluid modeling to a material response code to model in-depth surface conduction and accounts for space-charge-limited emission. The effectiveness of ETC for multiple test cases are investigated including a case with a sharp leading edge, case with in-depth material conduction, and a blunt body (i.e. capsule). For each of these test cases, ETC affects the surface properties, most notably the surface temperature, suggesting that ETC occurs for bodies in thermally intense, ionized flows, no matter the shape of the leading edge. An equation is provided to estimate the heat transfer induced by ETC.},
 bibtype = {inproceedings},
 author = {Hanquist, Kyle M. and Boyd, Iain D.},
 doi = {10.2514/6.2018-1714},
 booktitle = {AIAA Aerospace Sciences Meeting, 2018},
 keywords = {etc}
}

Electron transpiration cooling (ETC) is a proposed thermal management approach for the leading edges of hypersonic vehicles that utilizes thermionic emission to emit electrons to carry heat away from the surface. This paper presents a modeling approach for implementing ETC in a computational fluid dynamics (CFD) framework and assesses the modeling approach using a set of previously completed experiments. The modeling approach includes coupling the fluid modeling to a material response code to model in-depth surface conduction and accounts for space-charge-limited emission. The effectiveness of ETC for multiple test cases are investigated including a case with a sharp leading edge, case with in-depth material conduction, and a blunt body (i.e. capsule). For each of these test cases, ETC affects the surface properties, most notably the surface temperature, suggesting that ETC occurs for bodies in thermally intense, ionized flows, no matter the shape of the leading edge. An equation is provided to estimate the heat transfer induced by ETC.

@inproceedings{
 title = {Aerothermodynamic Design Optimization of Hypersonic Vehicles},
 type = {inproceedings},
 year = {2018},
 publisher = {AIAA Paper 2018-3108},
 city = {Atlanta, GA},
 id = {3d38af90-793f-3b9e-aee3-e39dae5583ca},
 created = {2021-01-05T04:56:35.848Z},
 accessed = {2021-01-04},
 file_attached = {true},
 profile_id = {6476e386-2170-33cc-8f65-4c12ee0052f0},
 last_modified = {2021-01-05T05:45:11.514Z},
 read = {false},
 starred = {false},
 authored = {true},
 confirmed = {false},
 hidden = {false},
 citation_key = {eyi:aviation:2018},
 source_type = {inproceedings},
 private_publication = {false},
 abstract = {The objective of this study is to develop a reliable and efficient design optimization method for hypersonic vehicles focused on aerothermodynamics. Considering the nature of hypersonic flight, a high-fidelity aerothermodynamic analysis code is utilized for the simulation of weakly ionized hypersonic flows in thermo-chemical non-equilibrium. A gradient-based method is implemented for optimization. Bezier or NURBS curves are used to parametrize the geometry or the geometry change. Linear elasticity theory is implemented for mesh deformation. Penalty functions are utilized to prevent undesired geometrical changes. The design objective is to minimize drag without increasing the heat transfer rate and the maximum values of the surface heat flux, temperature and pressure. Design optimizations are performed at different trajectory points of the IRV-2 vehicle. The effects of parametrizations, the number of design variables and freestream conditions on design performance are studied.},
 bibtype = {inproceedings},
 author = {Eyi, Sinan and Hanquist, Kyle M. and Boyd, Iain D.},
 doi = {10.2514/6.2018-3108},
 booktitle = {2018 Multidisciplinary Analysis and Optimization Conference}
}

The objective of this study is to develop a reliable and efficient design optimization method for hypersonic vehicles focused on aerothermodynamics. Considering the nature of hypersonic flight, a high-fidelity aerothermodynamic analysis code is utilized for the simulation of weakly ionized hypersonic flows in thermo-chemical non-equilibrium. A gradient-based method is implemented for optimization. Bezier or NURBS curves are used to parametrize the geometry or the geometry change. Linear elasticity theory is implemented for mesh deformation. Penalty functions are utilized to prevent undesired geometrical changes. The design objective is to minimize drag without increasing the heat transfer rate and the maximum values of the surface heat flux, temperature and pressure. Design optimizations are performed at different trajectory points of the IRV-2 vehicle. The effects of parametrizations, the number of design variables and freestream conditions on design performance are studied.

@phdthesis{
 title = {Modeling of Electron Transpiration Cooling for Leading Edges of Hypersonic Vehicles},
 type = {phdthesis},
 year = {2017},
 keywords = {etc},
 websites = {http://hdl.handle.net/2027.42/138537},
 city = {Ann Arbor},
 institution = {University of Michigan},
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@article{
 title = {Detailed modeling of electron emission for transpiration cooling of hypersonic vehicles},
 type = {article},
 year = {2017},
 keywords = {etc},
 pages = {1-13},
 volume = {121},
 publisher = {American Institute of Physics Inc.},
 id = {444e5e91-a821-3d5f-8328-b5515f1eb6b5},
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 abstract = {Electron transpiration cooling (ETC) is a recently proposed approach to manage the high heating loads experienced at the sharp leading edges of hypersonic vehicles. Computational fluid dynamics (CFD) can be used to investigate the feasibility of ETC in a hypersonic environment. A modeling approach is presented for ETC, which includes developing the boundary conditions for electron emission from the surface, accounting for the space-charge limit effects of the near-wall plasma sheath. The space-charge limit models are assessed using 1D direct-kinetic plasma sheath simulations, taking into account the thermionically emitted electrons from the surface. The simulations agree well with the space-charge limit theory proposed by Takamura et al. for emitted electrons with a finite temperature, especially at low values of wall bias, which validates the use of the theoretical model for the hypersonic CFD code. The CFD code with the analytical sheath models is then used for a test case typical of a leading edge radius in a hypersonic flight environment. The CFD results show that ETC can lower the surface temperature of sharp leading edges of hypersonic vehicles, especially at higher velocities, due to the increase in ionized species enabling higher electron heat extraction from the surface. The CFD results also show that space-charge limit effects can limit the ETC reduction of surface temperatures, in comparison to thermionic emission assuming no effects of the electric field within the sheath.},
 bibtype = {article},
 author = {Hanquist, Kyle M. and Hara, Kentaro and Boyd, Iain D.},
 doi = {10.1063/1.4974961},
 journal = {Journal of Applied Physics},
 number = {5}
}

Electron transpiration cooling (ETC) is a recently proposed approach to manage the high heating loads experienced at the sharp leading edges of hypersonic vehicles. Computational fluid dynamics (CFD) can be used to investigate the feasibility of ETC in a hypersonic environment. A modeling approach is presented for ETC, which includes developing the boundary conditions for electron emission from the surface, accounting for the space-charge limit effects of the near-wall plasma sheath. The space-charge limit models are assessed using 1D direct-kinetic plasma sheath simulations, taking into account the thermionically emitted electrons from the surface. The simulations agree well with the space-charge limit theory proposed by Takamura et al. for emitted electrons with a finite temperature, especially at low values of wall bias, which validates the use of the theoretical model for the hypersonic CFD code. The CFD code with the analytical sheath models is then used for a test case typical of a leading edge radius in a hypersonic flight environment. The CFD results show that ETC can lower the surface temperature of sharp leading edges of hypersonic vehicles, especially at higher velocities, due to the increase in ionized species enabling higher electron heat extraction from the surface. The CFD results also show that space-charge limit effects can limit the ETC reduction of surface temperatures, in comparison to thermionic emission assuming no effects of the electric field within the sheath.

@article{
 title = {Evaluation of Computational Modeling of Electron Transpiration Cooling at High Enthalpies},
 type = {article},
 year = {2017},
 pages = {283-293},
 volume = {31},
 websites = {https://arc.aiaa.org/doi/10.2514/1.T4932},
 month = {4},
 publisher = {American Institute of Aeronautics and Astronautics Inc.},
 day = {26},
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 abstract = {Amodeling approach for electron transpiration cooling of high-enthalpy flight is evaluated through comparison to a set of experiments performed in a plasma arc tunnel for air and argon. The comparisons include air and argon flow at high enthalpies (27.9 and 11.6 MJ/kg, respectively), with a Mach number of 2.5 to 3. The conversion of the reported enthalpies and Mach numbers to freestream temperatures and velocities is discussed. The numerical approach is described, including implementation of a thermionic emission boundary condition and an electric field model. Also described is the implementation of a finite-rate chemistry model for argon ionization. Materials with different electron emission properties are also investigated, including graphite and tungsten. The comparisons include two different geometries with different leading-edge radii. The numerical results produce a wide range of emitted current due to the uncertainties in freestream conditions and emissive material properties, but they still agree well with the experimental measurements.},
 bibtype = {article},
 author = {Hanquist, Kyle M. and Alkandry, Hicham and Boyd, Iain D.},
 doi = {10.2514/1.T4932},
 journal = {Journal of Thermophysics and Heat Transfer},
 number = {2},
 keywords = {etc}
}

Amodeling approach for electron transpiration cooling of high-enthalpy flight is evaluated through comparison to a set of experiments performed in a plasma arc tunnel for air and argon. The comparisons include air and argon flow at high enthalpies (27.9 and 11.6 MJ/kg, respectively), with a Mach number of 2.5 to 3. The conversion of the reported enthalpies and Mach numbers to freestream temperatures and velocities is discussed. The numerical approach is described, including implementation of a thermionic emission boundary condition and an electric field model. Also described is the implementation of a finite-rate chemistry model for argon ionization. Materials with different electron emission properties are also investigated, including graphite and tungsten. The comparisons include two different geometries with different leading-edge radii. The numerical results produce a wide range of emitted current due to the uncertainties in freestream conditions and emissive material properties, but they still agree well with the experimental measurements.

@inproceedings{
 title = {Computational analysis of electron transpiration cooling for hypersonic vehicles},
 type = {inproceedings},
 year = {2017},
 pages = {1-12},
 websites = {https://arc.aiaa.org/doi/abs/10.2514/6.2017-0900},
 publisher = {AIAA Paper 2017-0900},
 city = {Grapevine, TX},
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 citation_key = {hanquist:scitech:2017},
 source_type = {inproceedings},
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 abstract = {Simulations of a leading edge of a hypersonic vehicle using computational fluid dynamics (CFD) and a material response code are presented in order to investigate the effect in-depth surface conduction has on electron transpiration cooling (ETC). ETC is a recently proposed thermal management approach. Previous numerical studies have shown that ETC can significantly lower the stagnation point surface temperature of sharp leading edges of hypersonic vehicles. However, these studies have neglected the effect of heat also being conducted into the material as opposed to only into the flow via radiative cooling and ETC. A modeling approach is presented for ETC, which includes the boundary conditions for electron emission from the surface, accounting for the electric field and space-charge limit effects within the near-wall plasma sheath. A material response code is used to determine typical values of in-depth surface conduction for the test cases studied. Since ETC materials are still being developed, a parametric study is conducted for a range of material properties pertinent to ETC. The results of this study are used to generate in-depth surface conduction profiles, which are implemented into the CFD framework. The CFD simulations show that including in-depth surface conduction results in lower surface temperatures than predicted with radiative and ETC cooling alone. This is because in-depth surface conduction complements radiative cooling and ETC by moving heat away from the surface, in the case of surface conduction by moving the energy into the material, allowing for a lower surface temperature. The results also show that ETC remains a major mode of heat transfer away from the surface, even with in-depth surface conduction. This suggests that ETC is still a promising mode of thermal management, especially since it transfers energy to the flow instead of into the material.},
 bibtype = {inproceedings},
 author = {Hanquist, Kyle M. and Boyd, Iain D.},
 doi = {10.2514/6.2017-0900},
 booktitle = {AIAA SciTech Forum - 55th AIAA Aerospace Sciences Meeting},
 keywords = {etc}
}

Simulations of a leading edge of a hypersonic vehicle using computational fluid dynamics (CFD) and a material response code are presented in order to investigate the effect in-depth surface conduction has on electron transpiration cooling (ETC). ETC is a recently proposed thermal management approach. Previous numerical studies have shown that ETC can significantly lower the stagnation point surface temperature of sharp leading edges of hypersonic vehicles. However, these studies have neglected the effect of heat also being conducted into the material as opposed to only into the flow via radiative cooling and ETC. A modeling approach is presented for ETC, which includes the boundary conditions for electron emission from the surface, accounting for the electric field and space-charge limit effects within the near-wall plasma sheath. A material response code is used to determine typical values of in-depth surface conduction for the test cases studied. Since ETC materials are still being developed, a parametric study is conducted for a range of material properties pertinent to ETC. The results of this study are used to generate in-depth surface conduction profiles, which are implemented into the CFD framework. The CFD simulations show that including in-depth surface conduction results in lower surface temperatures than predicted with radiative and ETC cooling alone. This is because in-depth surface conduction complements radiative cooling and ETC by moving heat away from the surface, in the case of surface conduction by moving the energy into the material, allowing for a lower surface temperature. The results also show that ETC remains a major mode of heat transfer away from the surface, even with in-depth surface conduction. This suggests that ETC is still a promising mode of thermal management, especially since it transfers energy to the flow instead of into the material.

@inproceedings{
 title = {Aerodynamic optimization of a golf driver using computational fluid dynamics},
 type = {inproceedings},
 year = {2017},
 pages = {1-8},
 websites = {https://arc.aiaa.org/doi/abs/10.2514/6.2017-0724},
 publisher = {AIAA Paper 2017-0724},
 city = {Grapevine, TX},
 id = {f5694a66-e8aa-3649-82c9-4e6ada90b003},
 created = {2021-01-05T04:42:32.200Z},
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 authored = {true},
 confirmed = {false},
 hidden = {false},
 citation_key = {neitzel:scitech:2017},
 source_type = {inproceedings},
 private_publication = {false},
 abstract = {Driving distance and accuracy are the two key characteristics to an ideal golf drive. Besides having correct swing mechanics, there are numerous approaches that have been advanced to improve driver distance and accuracy, including driver shape, size, and material throughout the history of golf. Currently, with strict equipment conformity regulations from the United States Golf Association (USGA), the shape of the golf driver is greatly bounded, resulting in designs with marked improvements in design performance becoming less common. The required blunt body shape of the golf driver leads itself to be highly affected by aerodynamic forces, specifically pressure and viscous drag. Although the general shape of the golf driver head is greatly defined, small changes in shape can affect the aerodynamics significantly. This paper focusing on using Navier-Stokes computational fluid dynamic (CFD) simulations to reduce the aerodynamic drag while also increasing the yaw stability of the golf driver. Results include a characterization of the flow field experienced during a golf swing as well as the drag analysis of a generic driver. The adjoint method is used to identify surfaces on the driver that are most sensitive to drag. Finally, an optimization approach is discussed to create a low-drag, stable driver with design constraints such as USGA conformity and other parameters important to driver design such as a low center-of-mass and high moment-of-inertia.},
 bibtype = {inproceedings},
 author = {Neitzel, Kevin J. and Hanquist, Kyle M.},
 doi = {10.2514/6.2017-0724},
 booktitle = {AIAA SciTech Forum - 55th AIAA Aerospace Sciences Meeting}
}

Driving distance and accuracy are the two key characteristics to an ideal golf drive. Besides having correct swing mechanics, there are numerous approaches that have been advanced to improve driver distance and accuracy, including driver shape, size, and material throughout the history of golf. Currently, with strict equipment conformity regulations from the United States Golf Association (USGA), the shape of the golf driver is greatly bounded, resulting in designs with marked improvements in design performance becoming less common. The required blunt body shape of the golf driver leads itself to be highly affected by aerodynamic forces, specifically pressure and viscous drag. Although the general shape of the golf driver head is greatly defined, small changes in shape can affect the aerodynamics significantly. This paper focusing on using Navier-Stokes computational fluid dynamic (CFD) simulations to reduce the aerodynamic drag while also increasing the yaw stability of the golf driver. Results include a characterization of the flow field experienced during a golf swing as well as the drag analysis of a generic driver. The adjoint method is used to identify surfaces on the driver that are most sensitive to drag. Finally, an optimization approach is discussed to create a low-drag, stable driver with design constraints such as USGA conformity and other parameters important to driver design such as a low center-of-mass and high moment-of-inertia.

@inproceedings{
 title = {Limits for thermionic emission from leading edges of hypersonic vehicles},
 type = {inproceedings},
 year = {2016},
 pages = {1-15},
 websites = {https://arc.aiaa.org/doi/abs/10.2514/6.2016-0507},
 publisher = {AIAA Paper 2016-0507},
 city = {San Diego, CA},
 id = {6fe77891-8a13-3a7d-a30f-8c0a031a1311},
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 starred = {true},
 authored = {true},
 confirmed = {false},
 hidden = {false},
 citation_key = {hanquist:scitech:2016},
 source_type = {inproceedings},
 private_publication = {false},
 abstract = {Simulations of electron transpiration cooling (ETC) on the leading edge of a hypersonic vehicle using computational fluid dynamics (CFD) are presented. The thermionic emission boundary condition and electric field model including forced diffusion are discussed. Different analytical models are used to describe the plasma sheath physics in order to avoid resolving the sheath in the computational domain. The first analytical model does not account for emission in the sheath model, so the emission is only limited by the surface temperature. The second approach models the emissive surface as electronically floated, which greatly limits the emission. The last analytical approach biases the emissive surface, which makes it possible to overcome space-charge limits. Each approach is compared and a parametric study is performed to understand the effects that the material work function, freestream velocity, and leading edge geometry has on the ETC effect. The numerical results reveal that modeling the sheath as a floated surface results in the emission, and thus ETC benefits, being greatly limited. However, if the surface is negatively biased, the results show that the emission can overcome space-charge limits and achieve the ideal ETC benefits predicted by temperature limited emission. The study also shows that, along with negatively biasing the surface, emission is enhanced by increasing the number of electrons in the external flowfield by increasing the freestream velocity.},
 bibtype = {inproceedings},
 author = {Hanquist, Kyle M. and Boyd, Iain D.},
 doi = {10.2514/6.2016-0507},
 booktitle = {54th AIAA Aerospace Sciences Meeting},
 keywords = {etc}
}

Simulations of electron transpiration cooling (ETC) on the leading edge of a hypersonic vehicle using computational fluid dynamics (CFD) are presented. The thermionic emission boundary condition and electric field model including forced diffusion are discussed. Different analytical models are used to describe the plasma sheath physics in order to avoid resolving the sheath in the computational domain. The first analytical model does not account for emission in the sheath model, so the emission is only limited by the surface temperature. The second approach models the emissive surface as electronically floated, which greatly limits the emission. The last analytical approach biases the emissive surface, which makes it possible to overcome space-charge limits. Each approach is compared and a parametric study is performed to understand the effects that the material work function, freestream velocity, and leading edge geometry has on the ETC effect. The numerical results reveal that modeling the sheath as a floated surface results in the emission, and thus ETC benefits, being greatly limited. However, if the surface is negatively biased, the results show that the emission can overcome space-charge limits and achieve the ideal ETC benefits predicted by temperature limited emission. The study also shows that, along with negatively biasing the surface, emission is enhanced by increasing the number of electrons in the external flowfield by increasing the freestream velocity.

@inproceedings{
 title = {Modeling of electron transpiration cooling for hypersonic vehicles},
 type = {inproceedings},
 year = {2016},
 pages = {1-12},
 websites = {https://arc.aiaa.org/doi/abs/10.2514/6.2016-4433},
 publisher = {AIAA Paper 2016-4433},
 city = {Washington, D.C.},
 id = {e4c3c2cf-2804-38e5-8fac-a7b5e9fb4823},
 created = {2021-01-05T04:40:39.065Z},
 accessed = {2021-01-04},
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 profile_id = {6476e386-2170-33cc-8f65-4c12ee0052f0},
 last_modified = {2021-01-05T05:45:11.350Z},
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 authored = {true},
 confirmed = {false},
 hidden = {false},
 citation_key = {hanquist:aviation:2016},
 source_type = {inproceedings},
 private_publication = {false},
 abstract = {Electron transpiration cooling (ETC) is a recently proposed approach to manage the high heating loads experienced at the sharp leading edges of hypersonic vehicles. Computational fluid dynamics can be used to investigate the feasibility of ETC in a hypersonic environment. A modeling approach is presented for ETC, which includes devloping the boundary conditions for electron emission from the surface, accounting for the electric field and space-charge limit effects within the near-wall plasma sheath. Two different analytical models for space-charge limited emission are discussed. The first model assumes that the electrons are emitted cold from the surface while in the second approach the emitted electrons have a finite temperature. The theory shows that emitted electrons with a finite temperature, referred to as warm emission in the present paper, can reach higher levels of emission. This is important because the benefit of ETC, mainly reduction in the surface temperature, is directly correlated to the level of electron emission from the surface. The space-charge limit models are assessed using 1D direct-kinetic plasma sheath simulations. The simulations agree well with the space-charge limit theory proposed by Takamura et al. for emitted electrons with a finite temperature. Both models are implemented into a CFD code, LeMANS, and run for a test case typical of a leading edge radius in a hypersonic flight environment. The CFD results show finite temperature theory results in a larger reduction in wall temperature because more electron emission is allowed for than the cold emission theory. However, even with the electrons being emitted with a finite temperature, the emission still reaches space-charge limits for the test case considered, which can limit the benefits of ETC.},
 bibtype = {inproceedings},
 author = {Hanquist, Kyle M. and Hara, Kentaro and Boyd, Iain D.},
 doi = {10.2514/6.2016-4433},
 booktitle = {46th AIAA Thermophysics Conference},
 keywords = {etc}
}

Electron transpiration cooling (ETC) is a recently proposed approach to manage the high heating loads experienced at the sharp leading edges of hypersonic vehicles. Computational fluid dynamics can be used to investigate the feasibility of ETC in a hypersonic environment. A modeling approach is presented for ETC, which includes devloping the boundary conditions for electron emission from the surface, accounting for the electric field and space-charge limit effects within the near-wall plasma sheath. Two different analytical models for space-charge limited emission are discussed. The first model assumes that the electrons are emitted cold from the surface while in the second approach the emitted electrons have a finite temperature. The theory shows that emitted electrons with a finite temperature, referred to as warm emission in the present paper, can reach higher levels of emission. This is important because the benefit of ETC, mainly reduction in the surface temperature, is directly correlated to the level of electron emission from the surface. The space-charge limit models are assessed using 1D direct-kinetic plasma sheath simulations. The simulations agree well with the space-charge limit theory proposed by Takamura et al. for emitted electrons with a finite temperature. Both models are implemented into a CFD code, LeMANS, and run for a test case typical of a leading edge radius in a hypersonic flight environment. The CFD results show finite temperature theory results in a larger reduction in wall temperature because more electron emission is allowed for than the cold emission theory. However, even with the electrons being emitted with a finite temperature, the emission still reaches space-charge limits for the test case considered, which can limit the benefits of ETC.

@inproceedings{
 title = {Comparisons of computations with experiments for electron transpiration cooling at high enthalpies},
 type = {inproceedings},
 year = {2015},
 pages = {1-13},
 websites = {https://arc.aiaa.org/doi/abs/10.2514/6.2015-2351},
 publisher = {AIAA Paper 2015-2351},
 city = {Dallas, TX},
 id = {64d84bad-5320-3fb5-91fa-30679bdcd008},
 created = {2021-01-05T04:28:38.314Z},
 accessed = {2021-01-04},
 file_attached = {false},
 profile_id = {6476e386-2170-33cc-8f65-4c12ee0052f0},
 last_modified = {2021-01-05T05:45:10.643Z},
 read = {false},
 starred = {true},
 authored = {true},
 confirmed = {false},
 hidden = {false},
 citation_key = {hanquist:aviation:2015},
 source_type = {inproceedings},
 private_publication = {false},
 abstract = {A modeling approach for electron transpiration cooling of high enthalpy ight is compared to a set of experiments performed in a plasma arc tunnel for nitrogen and argon. The comparisons include nitrogen and argon ow at high enthalpies, 12,000 btu/lb and 5,000 btu/lb respectively, with a Mach number of 2.5 to 3. Converting the provided enthalpies and Mach numbers to freestream temperatures and velocities is discussed. The numerical approach is described including implementation of a thermionic emission boundary condition. Also described is the implementation of a finite-rate chemistry model for argon ionization. Different emissive materials are also investigated including graphite and tungsten. The comparisons include two different geometries with different leading edge radii. The numerical results produce a wide range of emitted current due to the uncertainties in freestream conditions and emissive material properties, but still agree well with the experiments. Future work recommendations are provided that may improve the physical accuracy of the modeling capabilities used in the comparisons.},
 bibtype = {inproceedings},
 author = {Hanquist, Kyle M. and Boyd, Iain D.},
 doi = {10.2514/6.2015-2351},
 booktitle = {45th AIAA Thermophysics Conference},
 keywords = {etc}
}

A modeling approach for electron transpiration cooling of high enthalpy ight is compared to a set of experiments performed in a plasma arc tunnel for nitrogen and argon. The comparisons include nitrogen and argon ow at high enthalpies, 12,000 btu/lb and 5,000 btu/lb respectively, with a Mach number of 2.5 to 3. Converting the provided enthalpies and Mach numbers to freestream temperatures and velocities is discussed. The numerical approach is described including implementation of a thermionic emission boundary condition. Also described is the implementation of a finite-rate chemistry model for argon ionization. Different emissive materials are also investigated including graphite and tungsten. The comparisons include two different geometries with different leading edge radii. The numerical results produce a wide range of emitted current due to the uncertainties in freestream conditions and emissive material properties, but still agree well with the experiments. Future work recommendations are provided that may improve the physical accuracy of the modeling capabilities used in the comparisons.

@inproceedings{
 title = {Conceptual analysis of electron transpiration cooling for the leading edges of hypersonic vehicles},
 type = {inproceedings},
 year = {2014},
 websites = {https://arc.aiaa.org/doi/abs/10.2514/6.2014-2674},
 publisher = {AIAA Paper 2014-2674},
 city = {Atlanta, GA},
 id = {e3dadee8-e88b-3352-9e54-2af260b2e5f1},
 created = {2021-01-05T04:27:07.372Z},
 accessed = {2021-01-04},
 file_attached = {false},
 profile_id = {6476e386-2170-33cc-8f65-4c12ee0052f0},
 last_modified = {2021-01-05T05:45:10.665Z},
 read = {false},
 starred = {true},
 authored = {true},
 confirmed = {false},
 hidden = {false},
 citation_key = {alkandry:aviation:2014},
 source_type = {inproceedings},
 private_publication = {false},
 abstract = {Recent progress is presented in an ongoing effort to perform a conceptual analysis of possible electron transpiration cooling using thermo-electric materials at the leading edges of hypersonic vehicles. The implementation of a new boundary condition in the CFD code LeMANS to model the thermionic emission of electrons from the leading edges of hypersonic vehicles is described. A parametric study is performed to understand the effects of the material work function, the freestream velocity, and the leading edge geometry on this cooling effect. The numerical results reveal that lower material work functions, higher freestream velocities, and smaller leading edges can increase the cooling effect due to larger emission current densities. The numerical results also show that the electric field produced by the electron emission may not have a significant effect on the predicted properties. Future work recommendations are provided that may improve the physical accuracy of the modeling capabilities used in this study.},
 bibtype = {inproceedings},
 author = {Alkandry, Hicham and Hanquist, Kyle M. and Boyd, Iain D.},
 doi = {10.2514/6.2014-2674},
 booktitle = {AIAA AVIATION 2014 -11th AIAA/ASME Joint Thermophysics and Heat Transfer Conference},
 keywords = {etc}
}

Recent progress is presented in an ongoing effort to perform a conceptual analysis of possible electron transpiration cooling using thermo-electric materials at the leading edges of hypersonic vehicles. The implementation of a new boundary condition in the CFD code LeMANS to model the thermionic emission of electrons from the leading edges of hypersonic vehicles is described. A parametric study is performed to understand the effects of the material work function, the freestream velocity, and the leading edge geometry on this cooling effect. The numerical results reveal that lower material work functions, higher freestream velocities, and smaller leading edges can increase the cooling effect due to larger emission current densities. The numerical results also show that the electric field produced by the electron emission may not have a significant effect on the predicted properties. Future work recommendations are provided that may improve the physical accuracy of the modeling capabilities used in this study.