Characterization of accurate launch vehicle unsteady aerodynamics is critical for component and secondary structure vibroacoustic design. For the National Aeronautics and Space Administration's's (NASA) Space Launch System (SLS), aeroacoustic environments have been derived primarily through sub-scale wind tunnel testing. Both optical techniques and high frequency pressure measurements have been utilized across multiple testing facilities and numerous vehicle configurations to develop a range of preliminary and detailed environments. As the vehicle has matured and evolved, the data collected from each subsequent configuration has allowed for comparison studies which isolate the effects of certain outer mold line (OML) features on measured fluctuating pressure levels. This paper presents observations on some of those effects for features which include abort system protuberances, various fairings geometries, interstage flanges, and multibody interactions between a central core and fall away boosters. These features, and the flow conditions produced by them, are broadly applicable to many launch vehicle configurations.
Testing of identical dual-mode scramjet flowpath geometries in the freejet and direct-connect configurations was conducted in two separate facilities. These tests enabled a comparison study between the two configurations, which was directed toward the determination of the effects of inlet distortion and backpressure on the performance and operability of a dual-mode scramjet. Bulk flow conditions were matched between the two facilities at the isolator entrance plane to simulate Mach 4.8 flight, and a series of metrics were established to quantify similarities and differences. The effects of flowpath backpressure in the direct-connect case were seen to be isolated to regions close to the exhaust. An approximate 10% decrease in combustor pressure rise and consequently integrated pressure force were observed in the freejet configuration; shock train length, however, remained the same. Mode transition was delayed from an equivalence ratio of ∼0.5 in the direct-connect case to ∼0.7 in the freejet configuration. Further, ignition difficulties were experienced in the freejet tests which were not encountered in those of direct-connect tests, limiting the scope of comparable test points. This work represents the first attempt at a quantitative comparison in the literature of an identical direct-connect and freejet dual-mode scramjet. Nomenclature A = cross-sectional area D = isolator duct height F = integrated pressure force H = fuel injector ramp normal height L = shock train length M 0 = freestream Mach number M u = Mach number at shock train leading edge M c = Mach number at combustor entrance _ m TBIVmc = predicted freejet flowpath mass capture _ m TBIVnom = nominal freejet facility mass flow rate _ m test = as tested facility mass flow rate (direct connect or freejet) P s = static pressure P 0 = total pressure p d ∕p u = peak static pressure/static pressure at leading edge of shock train Re u = unit Reynolds number, m −1 Re θ = Reynolds number based on boundary-layer momentum thickness T s = static temperature T 0 = total temperature x = axial distance θ = boundary-layer momentum thickness at the leading edge of shock train ϕ = angle of flowpath divergence from the horizontal φ = fuel equivalence ratio
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