Expansion Tunnel Facility Research Articles

The fidelity of DSMC method to analyze the aerothermodynamics of the pure continuum flow regime

Abstract The statistical probabilistic Direct Simulation Monte Carlo (DSMC) method was developed to interrogate the flow regime, in which thermal non-equilibrium exists and flow properties are non-continuous. Unfortunately, the DSMC method had not been preferred for the pure continuum flow regime ascribed to the computational burden arises with the computation of a large number of simulated molecule trajectories. However, this study showed that the DSMC method can handle the number density of real molecules of the pure continuum flow regime. The DSMC predictions were in close proximity to the experimental measurements taken on an Orion aeroshell model at Mach 11.4, 12.4, and 12.5 of reacting air in an expansion tunnel facility. The measurements corresponded to laminar boundary layer conditions, whereby Schlieren images were taken to assess the shockwave encapsulating the Orion aeroshell. The DSMC model predicted no isolated bump of the stagnation point heat flux, which was in line with the experiments. At a 20° angle of attack of the aeroshell, the model could capture the non-axisymmetric flow-field, in which the shear layer was attached at the windward side aftbody and detached at the leeward side aftbody. Thinning of the shockwave with the increase of flow-field temperature and molecular dissociation was also visibly distinct. The outcomes suggest the fidelity of the DSMC method in capturing the molecular interactions and the associated reaction kinetics in regimes varying from continuum to a rarefied gas without any substantial compromise of accuracy.

Experimental and Numerical Investigation ofHypervelocity Carbon Dioxide Flow over Blunt Bodies

This paper represents ongoing efforts to study high-enthalpy carbon dioxide flows in anticipation of the upcoming Mars Science Laboratory and future missions. The work is motivated by observed anomalies between experimental and numerical studies in hypervelocity impulse facilities. In this study, experiments are conducted in the hypervelocity expansion tube that, by virtue of its flow acceleration process, exhibits minimal freestream dissociation in comparison with reflected shock tunnels, simplifying comparison with simulations. Shock shapes of the laboratory aeroshell at angles of attack of 0, 11, and 16 deg and spherical geometries are in very good agreement with simulations incorporating detailed thermochemical modeling. Laboratory shock shapes at a 0 deg of attack are also in good agreement with data from the LENS X expansion tunnel facility, confirming results are facility-independent for the same type of flow acceleration. The shock standoff distance is sensitive to the thermochemical state and is used as an experimental measurable for comparison with simulations and two different theoretical models. For low-density small-scale experiments, it is seen that models based upon assumptions of large binary scaling values do not match the experimental and numerical results. In an effort to address surface chemistry issues arising in high-enthalpy groundtest experiments, spherical stagnation point and aeroshell heat transfer distributions are also compared with the simulation. Heat transfer distributions over the aeroshell at the three angles of attack are in reasonable agreement with simulations, and the data fall within the noncatalytic and supercatalytic solutions.

Scramjet Experiments in an Expansion Tunnel: Evaluated Using a Quasi-Steady Analysis Technique

As scramjet engine development moves toward higher and higher speeds, it will be important to continue to ground test complete nose-to-tail configurations at true flight conditions. Above Mach 10, the freestream total pressure requirements can only be met by an expansion-tube facility. To establish the practicality of ground-testing complete scramjet configurations at high Mach numbers, an expansion-tunnel facility is used to test a generic two-dimensional scramjet at a Mach 10 replication condition. As the flow produced in an expansion-tube facility is transient in nature, a method for analyzing point measurements within the scramjet is developed. A transient numerical simulation of the engine configuration without fuel injection is used to verify the validity of the technique, as well as to provide a reference data set for comparison to experimental data. These experiments show that stable supersonic combustion is established during the test time. Using the developed transient analysis, accurate quasi-steady pressure data at true flight conditions can be inferred from scramjet test measurements in an expansion-tube facility.