Effect of Thermally Induced Deformation in UVa Supersonic Combustion Facility

Journal of Propulsion and Power

et al. 2015

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

Section: Resultssupporting

confidence: 68%

Comparison of a Direct-Connect and Freejet Dual-Mode Scramjet

Steva

Journal of Propulsion and Power

et al. 2015

“…Furthermore, many experiments rely on a limited number of flow diagnostic methods that give an incomplete basis for evaluating numerical results. The scramjet flowpath at the University of Virginia Supersonic Combustion Facility (UVaSCF) has been used as a basis for several recent CFD validation studies that relied exclusively on measured wall pressures as the means of comparison [2][3][4][5]. Highly detailed experimental data sets suitable for more complete model validations have generally been limited to simple flows, such as the OSD/TRMC Test Media Effects coaxial supersonic combusting jet [6,7], which lack many characteristics of a scramjet combustor.…”

Section: Introductionmentioning

confidence: 99%

Collaborative Experimental and Computational Study of a Dual-Mode Scramjet Combustor

Journal of Propulsion and Power

Rice

et al. 2014

Advanced computational models of hypersonic air-breathing combustion processes are being developed to better understand and predict the complex flows within a dual-mode scramjet combustor. However, the accuracy of these models can only be quantified through comparison to experimental databases. Moreover, the quality of computational results is dependent on accurate and detailed knowledge of the combustor inflow and boundary conditions. Toward these ends, this paper describes results from a collaboration of experimental and computational investigators. Detailed computational fluid dynamics and finite element analyses were performed throughout the design and implementation of experiments involving a direct-connect scramjet combustor operating at steady state during long duration testing. The test section hardware was designed to provide substantial access for optical laser diagnostics. Measurement locations included the inflow plane and several locations downstream of fuel injection. A suite of advanced in-stream diagnostics were applied, many of which are described in companion papers. Significant results in this paper include measured static wall pressures and temperatures, stereoscopic particle image velocimetry, and focused schlieren imaging. Validated thermal finite element calculations in the scramjet hardware and temperature maps of the flow path boundaries are also presented. Comparison of experimental results with computational fluid dynamics predictions are discussed in a separate paper. Nomenclature H = fuel injector normal ramp height h = convective heat transfer coefficient k = thermal conductivity P ref = measured static pressure immediately downstream of facility nozzle P o = flow stagnation pressure _ q = rate of heat transfer T o = flow stagnation temperature T r = recovery or adiabatic wall temperature T w = wall surface temperature x = axial coordinate γ = specific heat ratio φ = fuel equivalence ratio

“…The scramjet flowpath at the University of Virginia Supersonic Combustion Facility (UVaSCF) has been used as a basis for several recent CFD validation studies that relied exclusively on measured wall pressures as the means of comparison. [2][3][4][5] Highly detailed experimental data sets suitable for more complete model validations have generally been limited to simple flows, such as the OSD/TRMC Test Media Effects coaxial supersonic combusting jet, that lack many characteristics of a scramjet combustor. 6,7 The current state of the art data set for supersonic scramjet combustion is due to the SCHOLAR experiment, which evaluated temperature and flow composition in a representative scramjet combustor using the dual-pump Coherent Anti-Stokes Raman Scattering (CARS) technique.…”

Section: Introductionmentioning

confidence: 99%

Close-Collaborative Experimental and Computational Study of a Dual-Mode Scramjet Combustor

50th AIAA Aerospace Sciences Meeting Including the New Horizons Forum and Aerospace Exposition

Rice

et al. 2012

Advanced computational models of hypersonic air-breathing combustion processes are being developed to better understand and predict the complex flows within a dual-mode scramjet combustor. However, the accuracy of these models can only be quantified through comparison to experimental databases. Moreover, the quality of computational results is dependent on accurate and detailed knowledge of the combustor inflow and boundary conditions. Toward those ends, this paper describes the initial results of a unique, close collaboration of experimental and computational approaches. Detailed computational fluid dynamics (CFD) and finite element thermal-structural analyses (FEA) have been performed throughout the design and implementation of a direct-connect scramjet combustor operating at steady state during long duration testing on the order of an hour or more. The test-section hardware has been designed to provide numerous access points for optical laser diagnostic measurements. Measurement locations include the inflow plane to the scramjet combustor as well as several locations downstream of the fuel injector. In addition, static wall pressures and temperatures are measured at numerous points along the fuel injector side of the scramjet flowpath. Initial CFD calculations were used to generate detailed thermal boundary conditions that were then applied to a non-linear, thermalstructural finite element model of the test-section. The calculated temperatures and thermal deformations are evaluated and validated against experimental measurements. Significant results described in this paper include experimentally measured static wall pressure and temperature data, Stereoscopic Particle Image Velocimetry (SPIV) and focused schlieren imaging. Validated finite element calculations of temperature in the test-section hardware, and temperature maps of the flowpath boundaries are also presented. CFD results are discussed in a separate paper.