| description abstract | This experimental study characterizes the performance of a radial rotating detonation combustor (RDC). An aerospike nozzle for rocket propulsion has been integrated into the center of the combustor, although the same combustor could also be coupled with turbomachinery. The radial RDC (RRDC) utilized a rapid to gradual (RTG) area change in the flow direction to effectively confine the detonation region close to the inlet plane and to improve the uniformity of the flow exiting the RDC. Three test cases were analyzed, (a) a baseline case at a total reactant mass flowrate, m˙ = 0.136 kg/s and equivalence ratio, ϕ = 0.6, (b) a higher reactant flowrate, m˙ = 0.318 kg/s and ϕ = 0.6, and (c) a higher ϕ = 0.8 at m˙ = 0.318 kg/s. All tests were conducted using methane and a 67% oxygen and 33% nitrogen (by mole) oxidizer mixture. Measurements were acquired using CTAP probes inside the combustion channel and along the aerospike to characterize the performance, PCB and ion probes near the detonation region to identify wave modes and their variations during the test, and thrust measurements using a six-axis force sensor. Results show highly complex wave modes with multiple corotating and/or counter-rotating waves depending upon the reactant flowrate. The pressure and thrust measurements are consistent with the wave mode analysis. In general, a positive (combustor only) pressure gain was inferred when losses associated with the injection system were excluded. The study highlights the challenges associated with fuel–air mixing and integrating the RDC with downstream hardware. | |
