Hicham Alkandry

Evaluation of Finite-Rate Surface Chemistry Models for Simulation of the Stardust Reentry Capsule

in the numerical analysis of entry ows. The rst part of this paper describes the implementation of an existing nite-rate surface chemistry module into a Navier-Stokes solver. The module can simulate the chemical interaction of the hypersonic gas ow with the surface of a planetary entry vehicle, and allows the specication of several dierent surface reaction types, such as adsorption, Eley-Rideal recombination, Langmuir-Hinshelwood recombination, and sublimation. In the second part of the paper, the eects of using two dierent surface chemistry models developed by Driver et al. and Park on the numerical predictions of the ow eld and surface properties of the Stardust return capsule at the 81 km trajectory point are evaluated. Both models include carbon oxidation, and one of the main dierences between the two is that a mechanism for carbon nitridation is included in Park’s model. The production of CN in Park’s model causes the mass blowing rate due to the removal of bulk carbon to be greater by as much as 45% compared to Driver’s model. The heat transfer to the surface for Park’s model is approximately 20% less than for Driver’s model. When the carbon nitridation reaction is excluded from Park’s model, the results show better agreement in the heat ux and blowing rate as compared to Driver’s model.

Investigation of the Interactions of Reaction Control Systems with Mars Science Laboratory Aeroshell

Interactions between the Reaction Control System (RCS) jets and the bow shock from the aeroshell of a Mars Science Lab (MSL) model are investigated. Images are obtained experimentally at the University of Virginia using a low-density, hypersonic wind tunnel with the Planar Laser Induced Iodine Fluorescence technique. The models are .44% MSL aeroshells fitted with 0.5 mm RCS orifices to simulate Reaction Control Systems in both parallel and transverse jet directions relative to the aeroshells. Experiments are conducted at Mach 12 in the underexpanded jet freestream flowfield with sonic RCS jets. Images for both transverse and parallel jets are obtained for nozzle-thrust coefficients ranging from 0 to 3. It is found that there is much interaction between the aeroshell bow shock and the RCS jet for the transverse jet cases; however, there was not much interaction between the parallel jet and the bow shock on the aeroshell. Results from a nozzle-thrust coefficient of 0.5 were compared to numerical simulations for similar conditions obtained using CFD at the University of Michigan. It is found that there is good agreement in flowfield density between the experimental and numerical results in the jet core of the RCS but greater differences near the jet boundaries.

Comparison of Transport Properties Models for Flowfield Simulations of Ablative Heat Shields

The goal of this study is to evaluate the effects of different models for calculating the mixture transport properties on flowfield predictions of ablative heat shields. The Stardust sample return capsule at four different trajectory conditions is used as a representative environment for Earth entry. In the first part of the study, the results predicted using Wilke’s missing rule, with species viscosities calculated using Blottner’s curve fits and species thermal conductivities determined using Eucken’s relation, are compared to the results obtained using Gupta’s mixing rule with collision cross-section data. The heat transfer to the vehicle predicted using the Wilke/Blottner/Eucken model is found to be larger than the value obtained using the Gupta/collision cross-section model by as much as 60%. The Wilke/Blottner/Eucken model also produces a larger mass blowing rate due to the oxidation of bulk carbon by as much as 25% compared to the Gupta/collision cross-section model. In the second part of the study...

Interactions of Single-Nozzle Sonic Propulsive Deceleration Jets on Mars Entry Aeroshells

The effects of the propulsive decelerator (PD) jet Mach number on the flowfield, surface, and aerodynamic properties of a Mars entry aeroshell are investigated in Mach 12 laminar fl ow of I 2-seeded N2 gas. This is achieved using the computational fluid dynamics (CFD) code LeMANS, as well as the planar laser-induced iodine fluorescence (PLIIF) experimental technique. The results show that the flowfield features, such as the standoff distance of the bow and jet shocks, are all affected by the PD jet Mach number. The results also show that as the thrust coefficient increases, the flow around the aeroshell approaches a jet-only, no freestream configuration due to a PD jet shield. Therefore, the effects of the PD jet Mach number on the surface properties and the drag coefficient increases. As a result, the difference in the drag coefficient between the supersonic and sonic jets increases to as much as 25%. However, since the drag is inversely proportional to the nozzle thrust, the total axial forces for the supersonic and sonic jets are in close agreement, with a maximum difference of 4%. This result indicates that the overall deceleration performance of the aeroshell is only slightly affected by the PD jet Mach number for these particular conditions. The study also shows that propulsive deceleration with central PD jets may only be beneficial for thrust coefficients greater than 1.5 for both sonic and supersonic jets; a result that appears to be independent of the jet exit Mach number. Finally, qualitative comparisons between LeMANS and PLIIF show overall good agreement in the bow shock profile and standoff distance.

Numerical study of hypersonic wind tunnel experiments for Mars entry aeroshells

The combination of landing future high mass systems with small landing footprints on Mars may require the use of both propulsive deceleration (PD) and reaction control system (RCS) thrusters. However, the interactions between these jets and the supersonic or hypersonic freestream involve complex flow phenomena that are still not well understood. This paper describes numerical and experimental techniques that are used in an effort to develop physically accurate methods to compute these complex flow interactions. The paper also presents a numerical parametric study that is conducted using the computational fluid dynamics (CFD) code LeMANS. This study examines the effects of low temperature and density values and radially nonuniform freestream conditions in the hypersonic wind tunnel facility using an aeroshell based on the Mars Science Laboratory in Mach 12 flow of nitrogen gas with PD and RCS jets off. It is shown that although the Blottner and the Sutherland models compute different values of viscosity at low temperatures, the flowfield and surface properties predicted by LeMANS using these two models are in very close agreement. The study also shows that thermal nonequilibrium effects are negligible. The radial freestream nonuniformities, however, have considerable effects on the flowfield and surface properties. The nonuniform conditions change the temperature and density distributions in most of the computational domain, widen the bow shock around the aeroshell, increase continuum breakdown regions, and decrease the drag coefficient of the capsule compared to the uniform conditions. Finally, the paper presents qualitative experimental comparisons with the computed results which show overall good agreement.

Coupled Flow Field Simulations of Charring Ablators with Nonequilibrium Surface Chemistry

This paper describes the coupling of a Navier-Stokes solver to a material response code to simulate nonequilibrium gas-surface interactions. The Navier-Stokes solver used in this study is LeMANS, which is a three-dimensional computational uid dynamics code that can simulate hypersonic reacting ows including thermo-chemical nonequilibrium eects. The material response code employed in this study is MOPAR, which uses the one-dimensional control volume nite-element method to model heat conduction and pyrolysis gas behavior. This coupling is demonstrated using a test case based on the Stardust sample return capsule. Coupled simulations are performed at three dierent trajectory conditions. The eects of the pyrolysis gas chemistry are evaluated by assuming that the gas is either in chemical equilibrium or composed entirely of non-reacting phenol. The results show that the non-reacting pyrolysis gas assumption produces higher convective heat uxes, surface temperatures, and mass blowing rates. These eects are mainly due to the composition of the pyrolysis gas. The additional species produced by the pyrolysis gas in the chemical equilibrium case react with oxygen and nitrogen atoms in the gas-phase. This results in fewer atoms participating in the exothermic surface reactions, which reduces the heat transfer to the vehicle.

Propulsion Deceleration Studies Using Planar Laser-Induced Iodine Fluorescence and Computational Fluid Dynamics

Future high-mass spacecraft entering the thin Martian atmosphere will require additional means of deceleration prior to deploying supersonic parachutes. Propulsive deceleration is one technology that is being considered. The interaction of the spacecraft aerodynamics with the propulsion deceleration (PD) jets has been shown to cause a decrease in drag coefficient with increasing thrust coefficient, which is not desirable for deceleration. Planar LaserInduced Iodine Fluorescence (PLIIF) images showed a lifting of the vehicle bow shock away from the aeroshell. Flowfield calculations performed using a CFD code showed that this lifting was responsible for the decrease in drag with increasing PD jet thrust. With 4 PD jets located midway between the aeroshell centerline and shoulder, PLIIF images showed that the vehicle bow shock is maintained between the jets as the thrust coefficient is increased. CFD calculations established that this bow shock was responsible for greater drag preservation with the peripheral jets. The peripheral jet drag coefficient was 4 times larger than the single jet value at a thrust coefficient of 2.0. The calculations also showed low pressure wakes located radially behind the peripheral jets which are responsible for the decrease in drag coefficient with increasing thrust coefficient and that high pressure is maintained between the jets. These results suggest that using a few peripheral PD jets located near the aeroshell shoulder would provide the greatest amount of drag preservation when using propulsive deceleration. Nomenclature

Conceptual Analysis of Electron Transpiration Cooling for the Leading Edges of Hypersonic Vehicles

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.

Comparison of Models for Mixture Transport Properties for Numerical Simulations of Ablative Heat-Shields

rst part of the study, the results predicted using Wilke’s mixing rule with species viscosities calculated using Blottner’s curve ts and species thermal conductivities determined using Eucken’s relation are compared to the results obtained using Gupta’s mixing rule with collision cross-section (CCS) data. The Wilke/Blottner/Eucken model overpredicts the heat transfer to the surface relative to the Gupta/CCS model by as much as 60%. The Wilke/Blottner/Eucken model also overpredicts the mass blowing rate due to the removal of bulk carbon by as much as 25% compared to the Gupta/CCS model. In the second part of the study, the eects of the mass diusion model are assessed using Fick’s, modied Fick’s, self-consistent eective binary diusion (SCEBD), and Stefan-Maxwell models. The results show that the oweld properties calculated using modied Fick’s, SCEBD, and Stefan-Maxwell models are in good agreement. Fick’s model overpredicts the heat transfer and mass blowing rate by as much as 20% relative to the Stefan-Maxwell model.

Investigations of Peripheral 4-Jet Sonic and Supersonic Propulsive Deceleration Jets on a Mars Science Laboratory Aeroshell

With the launch of Mars Science Laboratory (MSL), scheduled for 2011, Viking technology developed in the 1970s is reaching its limits for entry, descent and landing (EDL) on Mars, necessitating research and development of other technologies for decelerating high mass Mars entry systems (HMMES), such as propulsive deceleration (PD) jets. In this paper planar laser-induced iodine fluorescence is utilized to obtain qualitative flow visualization images and quantitative PD jet mole fraction images of peripheral sonic and supersonic PD jet models in Mach 12 flow and compared to CFD computations. The models are 0.22% of the MSL frontal area, with Mach 1 and Mach 2.66 jets on the frontal aeroshell of the model, oriented normal to the hypersonic flow. The interactions of PD jets with a Mach 12 freestream flow are visualized with coefficients of thrust (CT) varying from 0.5 to 3.0 in increments of 0.5. It was found that as CT increases the shock stand-off distance increases for both sonic and supersonic cases, with the supersonic distance at a CT = 3.0 being 17% greater than the sonic distance. The jet penetration distance was measured to be 50% greater for the supersonic case at a CT = 3.0. Experimental results were compared with CFD calculations of the sonic 4-jet configuration. Very good comparison was shown in the streamline patterns and jet mole fraction distributions. Using the validated CFD model, preliminary calculations showed that the drag coefficient for the 4-jet peripheral case was 3 times larger than that for the single centerline jet case at a CT of 0.5 and 6 times larger at a CT of 1.5, both with sonic exit conditions and the same total mass flow rate. The preservation of the vehicle drag was attributed to the normal bow shock between the peripheral jets which does not exist in the single centerline jet. The total axial force coefficient (sum of CT and CD) was calculated to be twice as large for the peripheral 4 sonic jets as for the single sonic centerline jet at a CT of 0.5 and 50% larger at a CT of 1.5. This result suggests that, for the same total mass flow rate and sonic exit Mach number, the propulsive deceleration performance of the peripheral 4-jet PD design will be considerably greater relative to the single centerline PD jet. This result is important for the design of PD jet decelerators for EDL for future HMMES missions.