Computer analysis of rarefied aerodynamics around a winged space-plane for Mars entry

Aerothermochemical analysis of the Orion capsule in rarefied transitional flow regime during Mars entry

It is well known that the Orion capsule was born as a rescue shuttle for the International Space Station. NASA developed the Orion project for the transportation of instruments and astronauts to moon, Mars, and beyond. Orion is similar in shape but larger than the Apollo command module; Orion will be able to host up to six astronauts, but Apollo command module could host only three astronauts. The aerodynamics of Orion during reentry to Earth in all rarefied regimes (that is, from free molecule regime to continuum) was already studied. The forthcoming use of Orion for Mars landing stimulated the author to evaluate global aerodynamic coefficients in rarefied flow along an entry path. The study considered problems both in axisymmetric and in three-dimensional flows by means of two direct simulation Monte Carlo codes. Global aerodynamic coefficients were computed at the altitudes of 80, 90, and 100 km. Special attention was paid to longitudinal stability.

Aero-thermo-dynamic analysis of a low ballistic coefficient deployable capsule in Earth re-entry

This paper is the follow-on of a previous paper by the author where it was pointed out that the forthcoming, manned exploration missions to Mars, by means of complex geometry spacecraft, involve the study of phenomena like shock wave-boundary layer interaction and shock wave-shock wave interaction also along the entry path in Mars atmosphere. The present paper focuses the chemical effects both in the shock layer and on the surface of a test body along the Mars orbital entry and compares these effects with those along the Earth orbital re-entry. As well known, the Mars atmosphere is almost made up of Carbon dioxide whose dissociation energy is even lower than that of Oxygen. Therefore, although the Mars entry is less energized than the Earth re-entry, one can expect that the effects of chemistry on aerodynamic quantities, both in the shock layer and on a test body surface, are different from those along the Earth re-entry. The study has been carried out computationally by means of a direct simulation Monte Carlo code, simulating the nose of an aero-space-plane and using, as free stream parameters, those along the Mars entry and Earth re-entry trajectories in the altitude interval 60-90 km. At each altitude, three chemical conditions have been considered: 1) gas non reactive and non-catalytic surface, 2) gas reactive and non-catalytic surface, 3) gas reactive and fully-catalytic surface. The results showed that the number of reactions, both in the flow and on the nose surface, is higher for Earth and, correspondingly, also the effects on the aerodynamic quantities.

The increasing interest in the exploration of Mars stimulated the authors to study aerodynamic problems linked to space vehicles. The aim of this paper is to evaluate the aerodynamic effects of a flapped wing in collaborating with parachutes and retro-rockets to reduce velocity and with thrusters to control the spacecraft attitude. 3-D computations on a preliminary configuration of a blunt-cylinder, provided with flapped fins, quantified the beneficial influence of the fins. The present paper is focused on Aerodynamics of a wing section (NACA-0010) provided with a trailing edge flap. The influence of the flap deflection was evaluated by the increments of aerodynamic force and leading edge pitching moment coefficients with respect to the coefficients in clean configuration. The study was carried out by means of two Direct Simulation Monte Carlo (DSMC) codes (DS2V/3V solving 2-D/3-D flow fields, respectively). A DSMC code is indispensable to simulate complex flow fields on a wing generated by Shock Wave-Shock Wave Interaction (SWSWI) due to the flap deflection. The flap angle has to be a compromise between the aerodynamic effectiveness and the increases of aerodynamic load and heat flux on the wing section lower surface.

A Computational Method to Optimize the Distribution of a Catalytically Active Material Inside a Nano-scale Pore

Mars Global Surveyor aerodynamics for maneuvers in Martian atmosphere

This study is an extension of former works by the present authors, in which the influence of the chemical models by Gupta and by Park was evaluated on thermo-fluid-dynamic parameters in the flow field, including transport coefficients, related characteristic numbers and heat flux on two current capsules (EXPERT and Orion) during the high altitude re-entry path. The results verified that the models, even computing different air compositions in the flow field, compute only slight different compositions on the capsule surface, therefore the difference in the heat flux is not very relevant. In the above mentioned studies, ionization was neglected because the velocities of the capsules (about 5000 m/s for EXPERT and about 7600 m/s for Orion) were not high enough to activate meaningful ionization. The aim of the present work is to evaluate the incidence of ionization, linked to the chemical models by Gupta and by Park, on both heat flux and thermo fluid-dynamic parameters. The present computer tests were carried out by a direct simulation Monte Carlo code (DS2V) in the velocity interval 7600-12000 m/s, considering only the Orion capsule at an altitude of 85 km. The results verified what already found namely when ionization is not considered, the chemical models compute only a slight different gas composition in the core of the shock wave and practically the same composition on the surface therefore the same heat flux. On the opposite, the results verified that when ionization is considered, the chemical models compute different compositions in the whole shock layer and on the surface therefore different heat flux. The analysis of the results relies on a qualitative and a quantitative evaluation of the effects of ionization on both chemical models. The main result of the study is that when ionization is taken into account, the Park model is more reactive than the Gupta model; consequently, the heat flux computed by Park is lower than the one computed by Gupta; using the Gupta model, in the design of a thermal protection system, is recommended.

The attitude aerodynamic control is an important subject in the design of an aerospace plane. Usually, at high altitudes, this control is fulfilled by thrusters so that the implementation of an aerodynamic control of the vehicle has the advantage of reducing the amount of thrusters fuel to be loaded on board. In the present paper, the efficiency of a wing-flap has been evaluated considering a NACA 0010 airfoil with a trailing edge flap of length equal to 35% of the chord. Computational tests have been carried out in hypersonic, rarefied flow by a direct simulation Monte Carlo code at the altitudes of 65 and 85 km, in the range of angle of attack 0-40 deg. and with flap deflection equal to 0, 15 and 30 deg.. Effects of the flap deflection have been quantified by the variations of the aerodynamic force and of the longitudinal moment. The shock wave-boundary layer interaction and the shock wave-shock wave interaction have been also considered. A possible interaction of the leading edge shock wave and of the shock wave arising from the vertex of the convex corner, produced on the lower surface of the airfoil when the flap is deflected, generates a shock wave whose intensity is stronger than those of the two interacting shock waves. This produces a consistent increment of pressure and heat flux on the lower surface of the flap, where a thermal protection system is required.

The attitude aerodynamic control is an important subject in the design of an aerospace plane. Usually, at high altitudes, this control is fulfilled by thrusters so that the implementation of an aerodynamic control of the vehicle has the advantage of reducing the amount of thrusters fuel to be loaded on board. In the present paper, the efficiency of a wing-flap has been evaluated considering a NACA 0010 airfoil with a trailing edge flap of length equal to 35% of the chord. Computational tests have been carried out in hypersonic, rarefied flow by a direct simulation Monte Carlo code at the altitudes of 65 and 85 km, in the range of angle of attack 0-40 deg. and with flap deflection equal to 0, 15 and 30 deg.. Effects of the flap deflection have been quantified by the variations of the aerodynamic force and of the longitudinal moment. The shock wave-boundary layer interaction and the shock wave-shock wave interaction have been also considered. A possible interaction of the leading edge shock wave and of the shock wave arising from the vertex of the convex corner, produced on the lower surface of the airfoil when the flap is deflected, generates a shock wave whose intensity is stronger than those of the two interacting shock waves. This produces a consistent increment of pressure and heat flux on the lower surface of the flap, where a thermal protection system is required.

The aim of the present thesis is the study of some problems of hypersonic rarefied Aerodynamics. More specifically the subjects here considered are: a) Development of a new parameter to detect non-equilibrium region. As already pointed out, the identification of non-equilibrium regions is important for an accurate solution of a flow field. More specifically an hybrid code needs a parameter to determine what method (DSMC or CFD), has to be used in the solution of the flow field. In this thesis, a new parameter to detect the non equilibrium region is proposed. This parameter is based on the Crocco theorem. The assumption on which the new parameter relies is that a theoretical relation, based on the hypotheses of equilibrium, as the one of the Crocco theorem, is not verified in non-equilibrium. The new parameter has been computed as the difference of the terms forming the Crocco theorem equation. Thus, one can expect that the higher is non-equilibrium, the larger is the mismatch between the terms and therefore the higher is the parameter. b) Improvement of approximate methods (bridging formulae) for the evaluation of aerodynamic coefficients of a re-entry vehicle in high altitude flight. At the first stage of a design of a re-entry vehicle it could be important to determine in a fast way the aerodynamic forces coefficient. The achieve this goal the well known “bridging formulae” are used. In this thesis a “new” methodology (here called “new” bridging formula) has been developed. The “new” bridging formula, has been successfully tuned to sphere and it has been also tested on two current capsules: EXPERT and ORION. c) Analysis and comparison of several chemical models: 1) peculiar of a DSMC approach such as quantic models (classic and new), Fan-Shen and Bird, 2) “classic” models such as the Gupta and Park models with and without ionization reactions. It is well known that one of the most important problems in the design of a capsule is the evaluation of heat flux during the re-entry. This evaluation has to provide information about the design of the Thermal Protection System (TPS). To this regard it is important to underline that due to the endothermic characteristic of the reactions, the chemical model affects the computation of the heat flux. A very deep analysis about the difference in the computation of heat flux between a direct simulation Monte Carlo code (DS2V) and a computational fluid dynamics code (H3NS) has been carried out. To this purpose a method to implement the Park model in a DSMC code has been developed. d) Application of DSMC codes to evaluate the aerodynamic coefficients of a current capsule (EXPERT) and a future aerospace vehicle (FAST20XX). For the EXPERT capsule computer tests have been carried out in the altitude flight with particular attention to the longitudinal stability of the capsule. As for as FAST20XX, the demonstration and validation of the numerical tools able to predict aero-thermal loads on a space re-entry vehicle at high altitude conditions has been carried out. This goal is fulfilled by the characterization of the DLR low density wind tunnel V2G and by an aerodynamic analysis of the available model of a lifting body.

The present study is the logical continuation of a former paper by the first author in which the influence of the chemical models by Gupta and by Park on the computation of heat flux on the Orion and EXPERT capsules was evaluated. Tests were carried out by the direct simulation Monte Carlo code DS2V and by the computational fluiddynamic (CFD) code H3NS. DS2V implements the Gupta model, while H3NS implements the Park model. In order to compare the effects of the chemical models, the Park model was implemented also in DS2V. The results showed that DS2V and H3NS compute a different composition both in the flow field and on the surface, even using the same chemical model (Park). Furthermore DS2V computes, by the two chemical models, different compositions in the flow field but the same composition on the surface, therefore the same heat flux. In the present study, in order to evaluate the influence of these chemical models also in a CFD code, the Gupta and the Park models have been implemented in FLUENT. Tests by DS2V and by FLUENT, have been carried out for the EXPERT capsule at the altitude of 70 km and with velocity of 5000 m/s. The capsule experiences a hypersonic, continuum low density regime. Due to the energy level of the flow, the vibration equation, lacking in the original version of FLUENT, has been implemented. The results of the heat flux computation verify that FLUENT is quite sensitive to the Gupta and to the Park chemical models. In fact, at the stagnation point, the percentage difference between the models is about 13%. On the opposite the DS2V results by the two models are practically equivalent.

A rarefied aero-thermo-dynamic analysis of a sample Earth Return Capsule during the high energy, high altitude re-entry path from an exploration mission is presented. The altitude interval 70-120 km is considered, where the capsule experiences different flow fields. In fact, the flow regime ranges from continuum low density to near free molecular flow and, even though the free stream velocity is almost constant (13 km/s) in the whole altitude interval, the Mach number changes from 44 to 32 and the Reynolds number, based on the capsule diameter, ranges from 4.92×104 to 9. The computations have been carried out using two direct simulation Monte Carlo codes: DS2V to compute local quantities such as heat flux, thermal and aerodynamic loads at zero angle of attack and DS3V to compute global aerodynamic coefficients in the range of the angle of attack 0-60 deg̤ The results verified that in this altitude interval the heat flux and the thermal load reasonably satisfy specific requirements for the thermal protection system and that the capsule is longitudinally stable up to an angle of attack of about 40 deg..

The computation of heat flux on two current re-entry capsules, European eXPErimental Reentry Testbed (EXPERT) and Orion, has been carried out by a direct simulation Monte Carlo code (DS2V) and by a computational fluid dynamic code (H3NS) in transitional regime, considering both non-reactive and fully catalytic surface. These capsules have been chosen for this analysis because they have been characterized by completely different shapes and re-entry trajectories. DS2V and H3NS use the Gupta and the Park chemical models, respectively. The results showed that the heat flux predicted by DS2V is always higher than that predicted by H3NS. Therefore, a sensitivity analysis of the chemical models on the heat flux has been carried out for both capsules. More specifically, the Park model has been implemented in DS2V as well. The results showed that DS2V and H3NS compute a different chemical composition both in the flow field and on the surface, even when using the same chemical model (Park); therefore, the different results obtained from the two codes can be attributed mostly to the different methodology used in handling all chemical processes.

Four different chemical models (Bird, Fan and Shen, ‘classic’, and ‘new’ quantic) have been implemented in a direct simulation Monte Carlo (DSMC) code for comparison purpose. Computer tests simulated a hypersonic, rarefied flow field past a bluff (or flat-ended) cylinder; the working gas was nitrogen. The flow field parameters were compared with those from the advanced DS2V code. The Stanton number at the stagnation point was compared also with the experimental data. The present analysis showed that the Fan—Shen model is comparable with the widely accepted Bird model. The results from the DSMC code based on these models are close to those computed by DS2V and to the experimental data. The good comparison of the Bird and the Fan—Shen models is a very important result both from a physical point of view, because the two models rely on different approaches, and from an operative point of view, because the Fan—Shen model could be considered as a possible alternative to the Bird model. On the contrary, the quantic models appear to be not suitable.