Reducing the weight and decreasing pressure losses of aviation gas turbine engines improves the thrust-to-weight ratio and improves efficiency. In ultra-compact combustors (UCC), engine length is reduced and pressure losses are decreased by merging a combustor with adjacent components using a systems engineering approach. High-pressure turbine inlet vanes can be placed in a combustor to form a UCC. In this work, experiments were performed to understand the performance and associated physics within a UCC. Experiments were performed using a combustor operating at pressures in the range of 520–1030 kPa (75–150 psia) and inlet temperature equal to 480–620 K (865 R–1120 R). The primary reaction zone is in a single trapped-vortex cavity where the equivalence ratio was varied from 0.7 to 1.8. Combustion efficiencies and NOx emissions were measured and exit temperature profiles were obtained for various air loadings, cavity equivalence ratios, and configurations with and without representative turbine inlet vanes. A combined diffuser-flameholder (CDF) was used to study the interaction of cavity and core flows. Discrete jets of air immediately above the cavity result in the highest combustion efficiencies. The air jets reinforce the vortex structure within the cavity, as confirmed through coherent structure velocimetry of high-speed images. The combustor exit temperature profile is peaked away from the cavity when a CDF is used. Testing of a CDF with vanes showed that combustion efficiencies greater than 99.5% are possible for 0.8 ≤ Φcavity ≤ 1.8. Temperature profiles at the exit of the UCC with vanes agreed within 10% of the average value. Exit-averaged emission indices of NOx ranged from 3.5 to 6.5 g/kgfuel for all test conditions. Increasing the air loading enabled greater mass flow rates of fuel with equivalent combustion efficiencies. This corresponds to increased vortex strength within the cavity due to the greater momentum of the air driver jets.

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