As part of a two-phase experimental study to obtain detailed heating and pressure data over the full-scale hypersonic international flight research and experimentation (HIFiRE-1, formally FRESH FX-1) flight geometry, Calspan-University at Buffalo Research Center has completed a matrix of ground tests to determine the optimal flight geometry and instrumentation configuration necessary to make measurements of desired flow phenomena during the flight experiment. The primary objective of the HIFiRE-1 flight experiment is to collect high-quality flight data from integrated flight instrumentation to be used for computational fluid dynamic code and ground test facility validation in regions of boundary-layer transition, turbulent separated flow, and shock/boundary-layer interaction. To support this flight experiment, data have been obtained in the large energy national shock hypervelocity wind tunnel employing a full-scale model over a range of Mach numbers from 6.5 to 7.4 and Reynolds numbers from 0:5E 06 to 5:5E 6 duplicating the reentry trajectory. These points gave researchers the best chance to measure the transition process on the forecone and have a turbulent separated flow on the cylinder that reattached onto the flare section. These test condition ranges were determined directly from the nominal descent trajectory of the Australian-launched Terrier-Orion launch vehicle that will serve as the booster for HIFiRE-1. The entire experimental database will be compared to future flight data and used by computationalists to validate codes in regions of attached and separated laminar and turbulent flows with shock/boundary layer interaction. In addition to the experimental data, Calspan-University at Buffalo Research Center performed a large amount of computational fluid dynamic analyses to confirm and validate not only the tunnel flow conditions, but also the two-and threedimensional flows over the model itself. These detailed computational results will be presented in a Part 2 companion paper.
A review is presented of ground test experiments of a 146-mm spherical capsule model with forebody and aftbody symmetry plane measurements of heating and pressure for a range of enthalpies and Reynolds numbers to obtain a dataset of fundamental validation data for CFD codes and to develop a database for design-of-experiment of future studies. Comparisons with laminar experiments are made using CFD demonstrating the influence of thermochemical non-equilibrium on the aerodynamic and aerothermal character of the body. For laminar flows in nitrogen up to 10MJ/kg, the good agreement with available measurements suggests that the description of the chemical and thermal activity of the gas is adequate. Analysis of the forebody heating found that a catalytic recombination probability of 0.002 to 0.010 was required to match the measured heat flux. For laminar flows in air up to 14 MJ/kg, some significant differences between CFD and measurements highlight the inadequacy of the current chemical and thermal models to predict the state of the gas after the rapid expansion in the nozzle. Analysis of forebody heating in these cases found that catalytic recombination probability near 1.0 was required to match measured heat flux, suggesting that potentially some type of excitation may be involved that is not properly modeled. Finally, the non-equilibrium excitation may be collision related as limited evidence suggests that the phenomenon becomes more benign as the number of collisions increases.
Predictions from the STABL code have been used to make comparisons to two series of fundamental transition experiments in a large-scale shock tunnel environment by solving the parabolized stability equations (PSE) to predict laminar-turbulent transition onset using a semi-empirical e N correlation. The two sets of experimental data were obtained at duplicated enthalpy Mach 10 conditions for slender geometries where transition is dominated by second-mode instability. The first experiment considered is a 7 O cone with sharp and blunt nosetips where the surface pressure gradient is zero and the second is an axisymmetric compression surface with a significant adverse pressure gradient acting upon the flow. The PSE analysis has predicted N-factor growth between 5.2 and 8.6 at the measured transition station for these cases, demonstrating a range of instability conditions describing the physical phenomena. The use of the Re θ /M E criterion is also explored, including examples on the axisymmetric compression surface where a low value indicating early transition shows the opposite trend to the more physically accurate PSE solution that indicates larger N-factor growth.American Institute of Aeronautics and Astronautics 37 TH AIAA Fluid Dynamics Conference and Exhibit AIAA 2007-4490 that of a flight condition. The freestream flow, which can be expanded to any test Mach number, is free of frozen, dissociated chemical contamination that can occur in the freestream of reflected shock facilities at very high enthalpies. CUBRC employs both reflected shock and expansion tunnels to provide a wide range of supersonic, hypersonic, and hypervelocity test capabilities, which are summarized in Fig 3. The LENS reflected shock-tunnel facilities were developed primarily to study the full-scale, hypervelocity flow physics of interceptors and air-breathing engine configurations. The scale and flow duplication capabilities of LENS are such that these vehicles can be studied at their full scale, inclusive of effects such as transition to turbulence, turbulent mixing from cross-flow jets and thrusters, duplicated flow chemistry, and other effects that are difficult or impossible to simulate at cold-flow or sub-scale conditions. Besides aerothermal measurements, extensive studies in this facility have been made using non-intrusive diagnostics such as aero-optic and aero-acoustic measurements, including recent work with tunable laser-diode diagnostics 3,4 . The capabilities of LENS-I duplicate the flight conditions of interceptors and scramjet engines from Mach 7 to 15 (with Reynolds number matching to Mach 22), while the LENS-II facility complements it in such a way that this capability is seamlessly extended down to Mach 3.5 at sea level density. III. Supporting Numerical Tools A. DPLR Navier-Stokes SolverAll ground test studies in the LENS facilities are extensively calibrated and validated with numerical tools. The primary CFD tool used is the DPLR code provided by NASA Ames Research Center. DPLR is a multi-block, structured, finite-vol...
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