View Video Presentation: https://doi.org/10.2514/6.2021-3685.vid Rotating Detonation Rocket Engines (RDREs) have the potential to increase performance and lower the cost of launch vehicles by harnessing the benefits of pressure gain combustion. Thermal management and modelling in RDREs is challenging due to the high heat transfer rates and complex heat transfer process. In this study, we made quantitative heat measurements in a laboratory RDRE operating at elevated chamber pressures (CTAP up to 173 psia) with GOX-GH2 and GOX-GCH4 propellants. We found the highest heat flux near the propellant injectors (up to 25 MW/m2). The heat flux near the injectors (detonation cell region) does not follow the mass flux scaling of the traditional Bartz equation. The heat flux in the downstream combustor region does follow the mass flux scaling of the traditional Bartz equation for heat flux in constant pressure combustion rocket engines.
View Video Presentation: https://doi.org/10.2514/6.2023-0355.vid Additively manufactured, water-cooled test hardware was developed for calorimetry analysis in a rotating detonation rocket engine architecture. The water-cooled chamber hardware was designed for extended-length hot-fire testing at the Air Force Research Laboratory (Rocket Propulsion Division) in Edwards, CA, and features sensors to recover local heat fluxes at locations of interest from the injector face and global heat transfer to the combustor walls. A design leveraging metal additive manufacturing was utilized to generate cooling channels to "snake" around sensor ports, enabling heavy instrumentation of a monolithic, axially cooled chamber without compromising coolant flow for traditional subtractive manufacturability. Conjugate heat transfer analyses (CHT) were performed to predict the temperatures within the chamber wall of the test hardware at steady state conditions for a given coolant flow rate and to determine the impact of channel snaking on local heat transfer and thermocouple placement. The calorimetry hardware developed is a critical step in demystifying the otherwise convoluted heat transfer mechanisms within rotating detonation rocket engine architectures, and will enable long-duration hot-fire tests hitherto unachievable with traditionally-manufactured rotating detonation rocket engine hardware designs. With the design complete, manufacture of the chamber is anticipated to be completed by the end of Winter 2023, and hot-fire testing and performance characterization of the chamber are planned to be performed at AFRL-Edwards in Spring 2023. Subsequent measurements of temperature and heat flux during hot-fire operation are expected to provide unique insights into the practical application of detonation-based thermodynamic cycles for rocket propulsion research and development.
Abstract : Triple Langmuir probes and emissive probes were used to measure the electron number density, electron temperature, and plasma potential downstream of a low-power Hall thruster. The results show a polytropic relation between electron temperature and electron number density throughout the sampled region. Over a large frat ion of the plume, the plasma potential obeys the predictions of ambipolar expansion. Near the thruster centerline, however, observations show larger gradients of plasma potential than can be accounted for by this means. Radial profiles of plasma potential in the very-near-field plume are shown to contain large gradients that correspond in location to the boundaries of a visually intense plasma region. [Copyright 2005 American Institute of Physics.]
Abstract : An experimental study of the measurement of cathode temperature, current distribution, and near-cathode electron number density in a high power hydrogen arcjet is presented. This study is motivated by the desire to better understand arc-electrode interactions in arcjet thrusters, which in many cases, is the main determinate of arcjet lifetime. Measurements such as these may also provide the needed boundary conditions for numerical arcjet simulations, presently under development. We describe in this paper the application of a non-intrusive in-situ measurement technique for on-axis, spectral imaging of the electrode region of arcjets, and the application of this technique to the measurement of the cathode and anode temperatures, cathode spot size, and current distribution in a 30kW hydrogen arcjet thruster. A relatively large field of view (twice the throat diameter) and high spatial resolution (9 micrometers) are achieved.
Abstract : A phase-bridge microwave interferometer operating at a frequency of 90 GHz (3mm wavelength) is currently under development at the Air Force Research Laboratory (AFRL) at Edwards AFB, California. The motivation for developing this diagnostic is the capability to take time resolved plasma density measurements with higher spatial resolution than other interferometers typically operating at 30 GHz. This interferometer has demonstrated preliminary electron density measurements in the plume of a 200 W Hall thruster. The interferometer has been modified to overcome initial difficulties encountered during the preliminary testing. The modifications include the ability to perform remote and automated calibrations as well as an aluminum enclosure to shield the interferometer from the Hall thruster plume. With these modifications, it will be possible to make unambiguous electron density measurements of the thruster plume as well as to rapidly and automatically calibrate the interferometer to eliminate the effects of signal drift. Due to the versatility of the diagnostic, it is also anticipated that it will be applied to Hall thrusters and large scale pulsed plasma sources under development.
Abstract : Preliminary results of the Air Force program investigating clustered Hall thrusters are presented, primarily experimental results on a cluster of four 200 W Busek BHT-200-X3 Hall thrusters. Preliminary measurements of plume current density, start transient interactions, cathode current sharing, and near exit plane magnetic fields are presented. Greatest thruster interaction occurs when cathodes are electrically connected. In a two thruster case, one cathode dominated electron emission, producing 90% of the required current. When the cathodes are electrically independent, the greatest cluster interaction occurs during a Start following exposure of the thruster discharge chambers to water vapor. In this case, the thrusters enter and exit a high anode current mode related to internal plasma oscillations in a non-continuous manner. This is unlike the typical smoothly continuous anode current transient of a single thruster. Individual thrusters appear able to affect the anode current mode, and presumably the plasma oscillations, of neighboring thrusters. Once the thrusters are conditioned and if the cluster is electrically unconnected, no significant interaction is observed. Plume ion current measurements of two thrusters have yielded what appears to he a slight narrowing of the ion current density profile from that e%pected from linear superposition of individual thruster measurements. Near exit plane magnetic field measurements indicate that the magnetic fields between the thrusters are affected by neighboring thruster magnetic fields. As such, the near plume electric fields would also he modified and may be responsible for apparent plume narrowing.
This work examines a rotating detonation rocket engine (RDRE) using large eddy simulations (LES) of a three-dimensional domain with discrete fuel and oxidizer injector orifices. The case follows an experimental set-up from the Air Force Research Laboratory (AFRL) with high-pressure gaseous CH4 and O2 as the reactants. Six high-fidelity computations are used to study the effects of varying equivalence ratio and mass flow rate as well as the addition of a throat at the end of the combustion chamber. The analysis first makes comparisons against available experimental results from qualitative and quantitative perspectives. In all instances, the simulations closely capture the macroscopic engine behaviors with minor deviations attributed to model choices. The focus then moves to examining the flow fields during stable operation, contrasting the simulations by means of detonation surfaces, spatially-averaged plots used for visualizing time-varying fields. From this reduced dimensionality, several trends are noted and observations can be made more readily. Increasing the mass flow rate over nominal conditions yields a scaled-up behavior containing nearly identical physical features but higher pressures, thrust, and Isp. In contrast, increasing the equivalence ratio leads to a divergence in the flow: substantially more waves are present in a dual-directional or slapping mode that dramatically alters the qualitative result, but fails to generate significant performance gain. Finally, the addition of a downstream constriction inevitably chokes the exiting mixture and adds thrust, while yielding counter-rotating waves and more parasitic deflagration.
