The direct approach of modeling the flow between all blade passages for each blade row in the compressor is too computationally intensive for practical design and analysis investigations with inlet distortion. Therefore a new simulation tool called the Turbine Engine Analysis Compressor Code (TEACC) has been developed. TEACC solves the compressible, time-dependent, 3D Euler equations modified to include turbomachinery source terms which represent the effect of the blades. The source terms are calculated for each blade row by the application of a streamline curvature code. TEACC was validated against experimental data from the transonic NASA rotor, Rotor 1B, for a clean inlet and for an inlet distortion produced by a 90-deg, one-per-revolution distortion screen. TEACC revealed that strong swirl produced by the rotor caused the compressor to increase in loading in the direction of rotor rotation through the distorted region and decrease in loading circumferentially away from the distorted region.

A small, uncooled plasma torch was developed and used in combination with a novel injector design to study ignition and flameholding in hydrogen-fueled supersonic flows. The plasma torch was operated with mixtures of hydrogen and argon at total flows of 4.7 x 10 3 to 3.30 x 10 4 seem. The fuel injector design consisted of five small upstream pilot fuel injectors, a rearward-facing step for recirculation, and three main fuel injectors downstream of the step. The plasma torch was located in the recirculation region and aU injection was perpendicular to the Mach 2 stream. The combustion tests were conducted in a direct-connect mode at simulated scramjet combustor conditions of Mach 2, total temperatures of 780-1560 K, and a static pressure of 101 kPa. Both semifreejet and ducted tests were conducted. The experimental results indicate that a low-power plasma torch operating on a 1:1 volumetric mixture of hydrogen and argon and located in the recirculation zone fueled by the upstream pilot fuel injectors is a good igniter for the flow conditions simulating a flight Mach number of 3.7 (T t , total temperature of the air = 780 K). The total temperature required to autoignite the hydrogen fuel for this injector geometry was 1470 K. The new injector configuration was shown to be a good flameholder over a wide range of total temperatures. Spectroscopic measurements were used to verify the presence of hydrogen atoms in the torch exhaust. Hydrogen atoms appear to be responsible for ignition at air total temperatures below 890 K. This observation is based on the results of the spectroscopic tests and the fact that a 100% argon plasma with the same power input did not ignite the flow at these temperatures.

A streamline curvature throughflow numerical approach is assessed and modified to better approximate the flow fields of transonic axial compression systems. Improvements in total pressure loss modeling are implemented, central to which is a physics-based shock model, to ensure accurate and reliable off-design performance prediction. The new model accounts for shock geometry changes, with shock loss estimated as a function of inlet relative Mach number, blade section loading (flow turning), solidity, leading edge radius, and suction surface profile. Data from a single-stage, isolated rotor provide the basis for experimental comparisons. Improved performance prediction is shown. The importance of properly accounting for shock geometry and loss changes with operating conditions is demonstrated.

Loss models used in compression system performance prediction codes are often developed from the study of two-dimensional cascades. In this paper, compressible fluid mechanics has been applied to the changes in shock geometry that are known to occur with back pressure for unstarted operation of supersonic compressor cascades. This physics-based engineering shock loss model is applicable to cascades with arbitrary airfoil shapes. Predictions from the present method have been compared to measurements and Navier–Stokes analyses of the LO30-4 and L030-6 cascades, and very good agreement was demonstrated for unstarted operation. A clear improvement has been demonstrated over previously published shock loss models for unstarted operation, both in the accuracy of the predictions and in the range of applicability. The dramatic increase in overall loss with increasing inlet flow angle is shown to be primarily the result of increased shock loss, and much of this increase is caused by the detached bow shock. For a given Mach number, the viscous profile loss is nearly constant over the entire unstarted operating range of the cascade, unless a shock-induced boundary layer separation occurs near stall. Shock loss is much more sensitive to inlet Mach number than is viscous profile loss.

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