Axial compressors are key components of aircraft engines, defining their operability. Active flow control is a powerful tool to optimize compressor stability, but the effect of these methods on the way the compressors run into stall remains unclear. In fact, previous works of the authors have shown that stall inception can be shifted from spike to modal type on the machine considered in this study. However, the transition from one mode to another is difficult to analyze with conventional methods. To explain the evolution of the stall process with an increasingly strong control action, the authors propose to study an experimental data set consisting of fast pressure transducers recordings measured during stall events with flow control. This work relies first on wavelet analysis but also proposes a new method inspired by system dynamic analysis to get a complementary view of the transition to stall for the controlled cases. The main finding is that spike instability disappears with only a very limited injected mass flow, as low as 0.4% of the main mass flow. Other interesting features are linked to the new method itself, providing a clear and computationally cheaper indicator of the stall process, and can be potentially investigated as stall warning method.
This work examines the catalytic ignition and self-sustained combustion of premixed fuel-rich methane–oxygen mixtures ([Formula: see text]). The experiments used a platinum microtube of 0.6 mm inner diameter with a continuous flow of reactants and subjected it to a linearly ramped power input until ignition was achieved. Following ignition, self-sustained combustion was established without external power input. The influence of the equivalence ratio and flow rate on the power input required to initiate combustion was obtained. Relatively less power was needed to ignite mixtures with increasing equivalence ratios and lower flow rates. Exhaust product analysis was performed during the self-sustaining combustion phase. Fourier transform infrared spectroscopy was employed to quantify the major combustion products, including carbon monoxide and carbon dioxide. Hydrogen production was also quantified using a thermal conductivity detector. Infrared imaging was used to determine the temperature of the catalytic microtube surface and track the postignition movement of the reaction front during the transition to steady combustion. The experimental results for the critical power input required to achieve ignition were compared with numerical simulations. A plug-flow assumption was used to model the microtube reactor, and surface and gas-phase reactions based on an existing kinetic model were included. The simulation results were used to obtain the ignition and hydrogen production maps for this microtube reactor.
Thermal management remains one of the critical challenges in the advancement of rotating detonation engine (RDE) technology for practical propulsion applications. To facilitate progress in this field, this study investigates the use of gaseous film cooling via axial injectors, aiming to characterize its influence on engine operation and cooling effectiveness. Experiments were conducted using a combination of gaseous oxygen and ethylene as propellants. This investigation comprises two stages: the initial phase employed an inert film coolant, specifically nitrogen, before transitioning to a fuel-based coolant utilizing ethylene. For each stage, a parametric study was conducted with coolant mass fractions of up to 23%. Across the 35 short-duration test fires conducted, stable wave propagation was sustained even with the introduction of film cooling. Performance degradation in specific impulse and characteristic velocity remained within acceptable limits of up to 29%, while heat flux was reduced by up to 34%. More importantly, the region surrounding the wave propagation zone, where heat loads were highest, was effectively reduced using this architecture. These findings demonstrate the potential of axially injected gaseous film cooling as an effective thermal management strategy for RDEs.
This work presents a combined experimental and computational investigation of the Cascaded Multi-Stage Impinging Jet (CAMUI) solid-fuel configuration for hybrid propulsion. CAMUI departs from conventional solid-fuel architectures by arranging orthogonal fuel blocks that force sequential oxidizer-jet impingement, thereby intensifying mixing and extending residence time, which could be of interest for solid-fuel ramjet and hybrid rocket applications. This study provides the first schematic and theoretical interpretation of CAMUI internal ballistics, combining firing experiments and computational fluid dynamics (CFD) simulations to explain its internal flowfield and the resulting postfiring fuel-grain geometry. Four oxygen/High-Density Polyethylene firing tests were performed while varying the oxidizer mass flow rate and interblock spacing. Combustion efficiencies exceeded 90%, with peak regression consistently occurring at the impingement point on each block’s front surface. CFD simulations reproduced internal velocity, temperature, and mixture-fraction fields, explaining the postfiring geometry of the front, back, and port surfaces. Front-surface regression was strongly nonuniform and governed by impinging temperature and port geometry. Back-surface regression was dominated by two counter-rotating vortices whose heat transfer intensified as the gap length decreased, while low-regression “hump” regions formed where the vortices merged and the flow separated from the wall. Inside the ports, regression was highest near the motor axis due to preferential oxidizer flow in the axis-adjacent region.
Modern gas turbines contain non-axisymmetric components (e.g., bifurcations and transition ducts) that disturb circumferential uniformity and drive long-range potential fields across blade rows, degrading performance. We introduce Turbomachinery Aerodynamic Vortical-Potential Integrated Solver (TAVIS), a reduced-order method formulated on curved streamsurfaces to capture component interactions at orders-of-magnitude lower cost than three-dimensional unsteady Reynolds-averaged Navier-Stokes (URANS). On each streamsurface, the flow is represented as a circumferentially uniform mean with potential/vortical perturbations. Velocity perturbations are decomposed into five harmonically resolved modes (intake, rotor-bound potential, rotor-bound circulation, rotor-shed vorticity, and stator circulation). Rotor effects are modeled using an actuator disc, where nonreflecting transmission is enforced through linearized continuity across the disc with the rotor Kutta condition. Combined with wall impermeability and the stator Kutta condition, all constraints are imposed simultaneously within a single finite-element linear system. Validation on high-bypass fan-outlet guide vane (OGV)-pylon configurations shows that TAVIS reproduces dominant harmonics and Mach-number distortions with band-aware RMSE<0.5% at mid- and tip-span; off-design discrepancies remain<4%. Solutions converge in ten seconds, 10(3) times faster than URANS, retaining accuracy. Including compressibility and geometric-variation effects, TAVIS delivers near-real-time predictions for optimization. The framework is readily extensible to other non-axisymmetric configurations (e.g., compressor struts, transition ducts, and propeller-rudder systems).
A rotating-detonation combustor (RDC) is subject to high heat loads attributable to high temperature gases and complex flow patterns. Ablator materials such as carbon-carbon composites suffer from erosion in a detonation region. Therefore, a gas-solid multiphase RDC (GSM RDC) system, which recovers heat loads by substituting the RDC wall with a solid fuel, is proposed. In this system, it is hypothesized that the detonation wave propagates as the solid fuel is consumed, causing the shock wave to interact with vaporized gas. This study investigated how the gasification of solid fuel affects detonation structure, combustion, and propulsion performance. Combustion tests are conducted in a cylindrical RDC with high density polyethylene (HDPE), oxygen, and ethylene. The combustion structure observed using a high-speed camera reveals that vaporized gas combusts along the chamber wall surface. The thrust in the RDC with the HDPE-wall configuration is 10% higher than that of the stainless steel (SS) wall. Propulsion performance can be predicted with +/- 4% accuracy by considering additional mass and modifying the equivalence ratio. These findings demonstrate the effect of vaporized gas combustion and help to clarify the interaction on a time-average basis. Thus, this combustor realizes a propulsion system for ablation cooling and thrust enhancement.
Aluminum agglomeration in solid rocket propellants has a significant impact on two-phase flow losses and combustion efficiency. Characterizing these agglomerates typically relies on high-speed photography, where manual analysis is labor-intensive and prone to subjectivity, while traditional image processing algorithms struggle with noise and motion blur. This study presents an automated pipeline for the detection of agglomerates using a Mask R-CNN deep learning architecture with a ResNet-50 backbone. The model was trained and evaluated on a dataset of 166 high-speed images containing over 8100 manually annotated instances, captured under varying pressure conditions (2-4 MPa). A shortest-chord diameter estimator is applied uniformly to all segmentation masks to reduce the systematic inflation that area- and perimeter-based estimators exhibit on motion-blurred masks. Detection performance reaches an F1-score of 0.82, and population-level agreement with the manual baseline yields a 1% difference in terms of D32 (Sauter mean diameter) and a 3% difference in terms of D43 (volume-weighted diameter). Our results demonstrate strong alignment with manual annotations, establishing deep learning as a highly scalable and effective alternative to manual analysis in this imaging regime.
Fuel sloughing, a phenomenon in which fragments of the fuel grain detach and eject from the chamber without burning, is a problem in paraffin-based hybrid rockets due to their low mechanical strength. Sloughing can cause nozzle blockage and overpressurization, thus risking motor failure. Embedding polymer lattices within paraffin wax has emerged as a promising strategy to enhance its strength and mitigate sloughing. However, the fundamental mechanisms that lead to sloughing and quantifying the mechanical strength enhancement due to the lattice remain limited. The current work aims to address these gaps through combined experimental and modeling efforts. First, the viscoelastic behavior of paraffin wax is characterized using rheological testing. Second, shear testing is performed on paraffin wax to evaluate the influence of lattice volume fractions on shear strength. Finally, a one-dimensional transient heat conduction model is developed to predict temperature distributions within the solid fuel grain during operation. The measured shear strengths are compared with aerodynamic shear stresses induced by oxidizer flow, establishing a criterion for the onset of sloughing. The combined results from the rheological, mechanical, and thermal analyses provide a framework for predicting sloughing, offering a pathway to improved design and reliability of paraffin-based hybrid rocket fuels.
The research group has previously proposed dimethyl ether (DME) as a promising alternative propellant for arcjet thrusters, given the toxicity and handling difficulties associated with hydrazine, which is widely used in current systems. However, the use of DME has been accompanied by the formation of soot on the cathode and nozzle surfaces during operation, leading to discharge instabilities that hinder reliable performance. To mitigate this issue, the researchers investigated the effect of adding nitrous oxide (N2O) to the DME propellant, aiming to promote oxidative reactions that suppress soot formation. Experimental results demonstrated that increasing the proportion of N2O in the mixture to a moderate extent significantly reduced soot deposition on the electrodes, thereby stabilizing the discharge and enhancing thrust performance. In particular, a DME/N2O mass ratio of 6:4 yielded the most favorable results, with no observable soot accumulation and without the excessive oxidation damage seen at higher N2O ratios on the cathode or nozzle, within the tested duration of 90 s. Reaction pathway analysis and experimental diagnostics revealed that reactive species, such as OH, O, and O-2, generated during the thermal decomposition of N2O, played a key role in suppressing soot formation.
This work characterizes a passively fed porous electrospray thruster firing Advanced Spacecraft Energetic Non-Toxic propellant (ASCENT, formerly AF-M315E), an energetic ionic liquid monopropellant. Results on thruster current-voltage characteristics, angular beam divergence, plume energy spectra, plume mass spectra, direct thrust measurements, and long-duration testing are reported. The propulsive performance is assessed via indirect and direct methods across multiple units. ASCENT is found to provide a thrust-to-power ratio of 40-65 mu N/W, a specific impulse of 600 s, and a total efficiency of 15%. The direct and indirect measurements of mass flow rate show significant disagreement, suggesting the presence of additional mass loss mechanisms. No failure modes, significant current decay, or degradation was observed over multiple thruster units up to 167 h firing time. These data can be used to advance the development of an ASCENT-fueled multimode chemical-electrospray propulsion system.
A vacuum arc thruster (VAT) exhibits significant potential for aerospace applications due to its robust structure and high specific impulse. The use of magnetic nozzles can enhance the performance. In this study, a grid-enhanced VAT with axial discharge was designed, and a magnetic field aligned with the discharge direction was applied for investigation. Results indicate that the electron temperature at the cathode spot and the plume region exhibits opposing trends. An electron cooling zone is observed approximately 40 mm downstream of the cathode, where the electron temperature decreases by up to 50%, corresponding to a 160.3% increase in plasma density and a 47% increase in ion velocity. Measurements of beam deposition reveal that the application of the magnetic field reduces lateral deposition by 77.2%, indicating that film replenishment primarily originates from low-velocity metal vapor and that a phenomenon of ion scattering through the grid was also observed. The rate of thrust increase diminishes with stronger magnetic fields, which is attributed to the suppression of the plume density by excessively strong magnetic fields, with an optimal magnetic field determined to be 190 G.
A major challenge of ramjet engines is their inability to produce appreciable thrust at subsonic speeds. The Atlantis Intake System (AIS) is a novel design for a ramjet inlet intended to produce thrust through a wide range of flight Mach numbers, including under static conditions. With the AIS, the ramjet operates as an ejector ramjet, using a supersonic hydrocarbon fuel jet to pressurize and entrain ambient air. A computational fluid dynamics study is performed to assess the change in performance of the AIS due to preheating the fuel jet using heat released from combustion. First, the computational model is assessed for two combustion-related problems relevant to an ejector ramjet: a three-dimensional bluff-body stabilized flame and a ramjet with the AIS operating under static conditions. The partially stirred reactor model is used for modeling turbulent combustion, and it is shown that using a modified mixing timescale with this model improves the accuracy of the computational results compared to the experiment. For an ejector ramjet, preheating the fuel jet is shown to increase the ratio of air to fuel entrained by the AIS and the specific impulse (Isp).
To use hybrid rockets as kick motors, it is essential to address their inherently low combustion efficiency issue. Although baffle plates are employed to improve combustion efficiency, their structural failure due to erosion phenomena caused by combustion gases becomes a serious concern. This study aims to identify the dominant parameters governing the erosion rate at the stagnation point of an impingement flow and to establish a predictive numerical model. Experiments were conducted using a high-density polyethylene/N2O hybrid rocket, varying the impinging jet velocity gradient and chamber pressure. A predictive model for the baffle plate erosion in hybrid rocket environments was developed by modifying the heterogeneous reaction mechanism in a preexisting theoretical model established in the field of carbon combustion. The experimental results revealed that the velocity gradient exponent was lower than the theoretical value under diffusion-limited conditions. Also, a dependence of the baffle plate erosion rate on the surface temperature was observed, indicating the erosion rate is not diffusion limited. The application of the theoretical model demonstrates good alignment between predicted and experimental results with an average error of about 33% without any corrections. This confirms the applicability of the theoretical model for predicting baffle plate erosion in hybrid rockets.
The inner surface pressure of an axisymmetric inlet/isolator model was measured using anodized-aluminum pressure-sensitive paint (PSP) viewing through cast acrylic. Temperature-sensitive paint was used to correct for the PSP's temperature sensitivity. The model was tested under Mach 5.7 flow at Re=7.1 & times;106/m under conventional noise conditions. Transverse jet injection with jet-to-inlet mass-flow ratios up to 0.8 was used to induce unstart in the inlet/isolator. Background-oriented schlieren visualization of the inlet shocks was collected simultaneously with the PSP to determine when the inlet unstarted. Computational fluid dynamics results were used to determine off-wall flow structures and quantify approach conditions in the isolator. The dominant frequency of the quasi-steady shock train was 500 to 600 Hz. Coherence analysis revealed a linear relationship between the unsteady pressures downstream of the shock-train leading edge (STLE). A phase difference of pi to 3 pi/2 was seen between the region immediately downstream of the STLE and the region over approximately 1.5 isolator diameters downstream of the STLE. The PSP and conventional transducer measurements of STLE pressure rise agreed well with each other. The unstart process took approximately 10 to 45 ms, and the peak velocities were 0.142 u infinity to 0.233 u infinity, depending on the jet-to-inlet mass-flow ratio. The quasi-steady STLE shock front shape was measured with high spatial resolution and matched that of an upstream inlet shock. The pressure profile shape changed quickly due to the presence of inlet shocks and transient changes due to mass-injection, which has implications to unstart detection methods.
In this paper, we present the design, characterization, and calibration of a counterbalanced pendulum thrust stand suitable for ultrasensitive measurements of periodically pulsed thrusters. High thrust sensitivity is obtained by implementing optical heterodyne detection of the thrust stand displacement, which we show achieves sub-nanometer resolution. In addition, we exploit the high quality factor of the thrust stand mechanical resonance to amplify the displacement response by several orders of magnitude. As a test of the thrust stand's high sensitivity, we perform a thrust calibration using light pressure generated by nine reflections of a continuous-wave laser with a power of less than 10 W. The calibration is performed at a resonant frequency of f=0.748 +/- 0.005 Hz and a quality factor of Q similar to 392, although quality factors as high as 9215 have been measured. Using a modulated square wave forcing with a 50% duty cycle, we measure a sensitivity slope of 464 +/- 20 nN/cycle and a detection limit of <100 nN. To the authors' knowledge, this is the first use of light pressure for counterbalanced pendulum thrust stand calibration and establishes a device sensitivity suitable for measuring modulated thrust levels in a variety of low-thrust applications.
This paper presents the development of algorithms and their implementation in a graphics processing unit (GPU)-accelerated computational fluid dynamics (CFD) solver for efficient steady and unsteady analyses of flowfields within multirow turbomachinery. To enhance convergence, the data-parallel lower-upper relaxation method is incorporated as a residual smoother within the Runge-Kutta framework. The method allows for a large time step while maintaining parallelism at a cell level. The mixing plane and sliding plane methods are implemented for multirow turbomachinery scenarios. Particular attention is paid to the implementation and optimization of these two rotor-stator coupling methods within the Compute Unified Device Architecture. Various strategies are employed to leverage the capabilities of GPUs for high-performance computation. The presented solver achieves speedup factors of about 18 and 24 for single- and double-precision floating-point arithmetic, respectively, in GPU computing under equivalent thermal power consumption to that in central processing unit computing. Parallel scalability tests on an eight-GPU cluster demonstrate high parallel efficiency, reaching 96% and 98% for strong and weak scalings, respectively. For validation, a series of steady analyses are conducted to analyze the flowfields within a fan stage and an axial compressor, demonstrating good agreement of the overall performance metrics and radial flow profiles with the corresponding experimental data. Furthermore, the stall behavior of the compressor at 65% design speed is analyzed, and the results are consistent with the experimental findings.
Mechanically processed aluminum powders containing small amounts of gallium or indium reduce the volume-mean diameter of molten agglomerations in an ammonium-perchlorate composite propellant by as much as 60%. In cases where very low concentrations of gallium or indium are used (3-5 wt %), increases in propellant burning rate were observed as well. Higher concentrations of gallium or indium (10-16 wt %) maintained smaller agglomeration sizes but did not alter the burning rate of the propellant. A mechanism explaining the tradeoff between burning rate and agglomeration potential was proposed: gallium and indium reduce the onset temperature of an aluminum particle, leading to quicker ignition, shorter surface residence time, and less time to agglomerate. However, high concentrations of gallium or indium significantly reduce flame temperatures, thereby leading to reduced heat feedback and lower burning rates. This work also discovered that high concentrations of gallium/indium reduce the maximum theoretical specific impulse of a motor by several seconds in some cases.
The instability of combustion and low rate of fuel regression of the traditional hybrid rocket motor limit its applications. The distributed tube injector (DTI) presents a promising solution by significantly enhancing average fuel regression rates up to 8 mm/s, while maintaining stable combustion, particularly at the laboratory scale. In this study, we investigate the DTI at high fluxes to determine the characteristics of stability of combustion in hybrid rockets. We conducted tests by using motors with paraffin-based fuel and nitrous oxide in the blowdown mode. They were configured to run at fluxes of around 800 kg/(m2 & sdot;s). Data from 13 hot-fire tests showed that pressure oscillations in the chamber did not exceed 0.5 bar, thus demonstrating that the DTI-configured hybrid motor could operate smoothly even at motor fluxes that were nearly four times higher than those of traditional hybrid motors. We also verified the capability of the DTI to improve the regression rate. The average regression rate of DTI-configured motors exceeded 6 mm/s, while classically configured hybrid motors achieved a rate of only about 1.5 mm/s at similar fluxes. The insights obtained from the tests can inform the design of hybrid motors that require a high thrust density.
