Thermodynamic coupling mechanisms in precooled turbine-based combined cycle: From fuel properties to multi-objective performance characterization

Conventional propellant materials, including polymers, metals, and doped polymers, have long been investigated for laser ablation propulsion, though achieving desired specific impulse values with low-intensity continuous wave laser irradiation remains a challenge. This study presents low-density graphene-delignified wood composites as novel, sustainable propellants for continuous-wave laser ablation propulsion. Experimental results show that natural wood achieves a density-specific specific impulse of 5043.0±188 s g-1 cm3, surpassing the performance of all previously reported conventional propellants. Graphene delignified wood improves the absorption of laser energy of natural wood by 98.61%, thereby achieving an ultra-low ablation threshold intensity of 0.54 MW m-2; the lowest intensity attained so far, and an exceptional density-specific specific impulse of 1569.60±57.40 s g-1 cm3. The absolute specific impulse values of NW and GDW are 907.74 and 800.49s, respectively, whereas no other material has achieved a value >100s under CW laser irradiation. Additionally, graphene delignified wood has a tensile strength of 273.1MPa, ∼10 times greater than NW, and a specific tensile strength of 533MPa g-1 cm3, exceeding common aerospace structural materials in strength-to-weight ratio. The findings unlock the potential of lightweight and sustainable graphene delignified wood for use in various light and space applications.

The design of solid rocket motors (SRMs) involves complex trade-offs between performance, structural integrity, and safety. This study presents an open-source computational framework that couples the openMotor simulation environment with a genetic algorithm (GA) to automate SRM geometry optimization. The framework enables exploration of multidimensional design spaces defined by user-specified constraints, targeting improved total impulse and thrust-time characteristics while maintaining structural safety margins. The algorithm evaluates simulation outputs using a weighted normalized penalty function. To validate the optimization results, two SRM prototypes—a baseline and an optimized design—were fabricated and tested on a horizontal static test stand. The optimized motor achieved an experimentally measured specific impulse corresponding to 89% of the simulated prediction. The results demonstrate that coupling open-source simulation with heuristic optimization produces realistic designs and reduces manual tuning. This framework establishes a foundation for adaptive optimization of SRMs using experimental feedback and open-source computational tools. • An open-source framework integrating openMotor with a genetic algorithm for automated SRM geometry optimization. • Demonstrated reduction of design–test cycles by coupling simulation and optimization with experimental calibration. • Experimental validation demonstrated agreement between optimized motor performance and simulation predictions.

Impulse scaling during magnetic reconnection, the magnetic energy conversion to kinetic energy, via direct Mach probe measurements in Magnetic Reconnection Experiment is examined. Ion exhaust velocity and impulse scalings with reconnecting magnetic field during the push phase of driven reconnection are presented. The outflows and impulse measurements are compared with global MHD simulations. Both measurements and simulations reveal a favorable scaling, greater than linear, of impulse with reconnecting field. These scaling results establish that magnetic reconnection could be utilized for plasma propulsion.

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.

Abstract This study investigates the effects of non-condensable gas (NCG) content (0-9%) on the steam direct contact condensation process in a large space using the VOF multiphase flow model and the Realizable k-e turbulence model. Phase change and heat and mass transfer are simulated by incorporating a UDF. The results demonstrate that NCG markedly suppresses steam condensation, leading to a transition in the morphology of the vapor plume from periodic contraction to stable ejection, accompanied by bubble detachment phenomena. As the NCG content increases, the axial penetration of the vapor plume is enhanced, and the flow field characteristics undergo significant changes. Downstream of the ejector, the fluctuations in temperature and velocity become more pronounced, while the pressure fluctuation frequency at the throat increases and the mass transfer rate decreases. The thrust performance, quantified by specific impulse, exhibits a non-monotonic trend with increasing NCG content: it first increases and then decreases, peaking at 7% NCG. This study provides a theoretical basis for the optimized design of steam ejectors and condensers under high NCG conditions.

Abstract Conventional modeling of two-stage wave-heating magnetoplasma thrusters faces fundamental limitations in capturing the intrinsically interconnected processes including plasma generation, transport, and wave heating. Thus, there is a huge gap between theoriodical studies and applications.Presented here is an integrated self-consistent model for the two-stage wave-heating magnetoplasma thruster by developing a multi-module framework iteratively linking specialized solvers for various processes important for thrust and specific impulse, which are crucial for space missions. This approach achieves the first self-consistent whole-device simulation of the 50-kW STAR thruster, resolving the complete process chain from helicon plasma generation to ion cyclotron acceleration.The results provide the three-dimensional distribution characteristics of plasma/neutral-gas density and electron/ion temperature, which are key parameters for aerospace thrusters. Model benchmark shows good agreement against STAR experiments. Furthermore, the correlations of plasma generation, transport, wave energy deposition, and their influences on plasma parameters are revealed, providing a foundation for thrust/specific-impulse throttling in deep-space missions.

Hybrid rocket engines offer a compromise between safety, controllability, and performance, making them attractive for small-scale propulsion systems. However, oxidizer selection remains a critical early-stage design decision that cannot be determined solely from ideal thermodynamic metrics. This study presents a comparative analysis of three oxidizers—nitrous oxide (N2O), gaseous oxygen (GOX), and liquid oxygen (LOX)—for a 1 kN-class hybrid rocket engine using HDPE fuel under identical operating conditions. Equilibrium combustion performance was first evaluated using NASA Chemical Equilibrium with Applications (CEA) to determine optimal oxidizer-to-fuel ratios and theoretical specific impulse. These results were subsequently refined using Rocket Propulsion Analysis (RPA) to incorporate finite combustion chamber geometry and non-ideal nozzle expansion effects. The equilibrium analysis predicts maximum specific impulses of approximately 260 s for N2O/HDPE and nearly 300 s for oxygen-based systems. However, finite-geometry modelling indicates that practical performance is reduced by approximately 5–8%, yielding delivered specific impulses of about 275 s for GOX and 272 s for LOX. The results demonstrate that although oxygen (GOX and LOX) provides higher thermodynamic performance, the practical advantage of LOX over GOX becomes marginal at the kilonewton scale. Consequently, oxidizer selection for small hybrid engines should be treated as a system-level trade-off involving performance, infrastructure complexity, and operational safety.

Utah State University has developed a high-performance “green” hybrid propulsion technology based on the unique electrical breakdown properties of 3D-printed acrylonitrile butadiene styrene. Using 3D-printed ABS as fuel, typical startup sequences require approximately 5–15 joules and, once started, the system can be sequentially fired with no additional energy inputs required. The number of possible ignitions is limited only by the amount of fuel. The most technologically mature version uses gaseous oxygen (GOX) as oxidizer and 3D-printed ABS as fuel. While GOX is mass-efficient, it lacks volumetric efficiency unless highly pressurized. Nytrox, a blend of GOX and nitrous oxide, improves propellant density and volumetric efficiency, while maintaining acceptable levels of mass efficiency (specific impulse). Nytrox can safely self-pressurize, eliminating the need for a separate oxidizer pressurization system and reducing overall complexity. However, employing Nytrox as a direct substitute for GOX results in reduced ignition reliability and considerably increases cold-start ignition latency. This paper quantifies the latency, explores its sources, and analyzes expected behaviors. Solutions include raising combustion and storage pressures to boost oxygen content in Nitrox’s liquid phase and increasing combustion chamber pressure to reduce ignition delays.

Abstract Microwave electrothermal thrusters enable electrodeless electrothermal propulsion with propellants beyond noble gases, but molecular propellants introduce chemistry that can compete with translational heating. We report an experimental comparison of hydrogen and ammonia in a 17.8-GHz resonant-cavity microwave electrothermal thruster at tens of watts. Because the present hardware employs a straight-orifice nozzle, chamber performance is evaluated via characteristic velocity ( $$c^*$$ ) calculated from hot-fire stagnation pressure and mass flow rate, with a nozzle-equivalent specific impulse estimate provided for scaling. Hydrogen demonstrates markedly higher chamber performance than ammonia. Ammonia, despite exhibiting a stronger stagnation-pressure increase, shows degraded characteristic velocity—evidence that absorbed power is preferentially partitioned into dissociation and internal excitation rather than equilibrated heating. Consistent with this interpretation, optical emission during ammonia operation shows NH/N 2 band features and Balmer H α emission. The results highlight a key design lesson for microwave electrothermal thrusters using molecular propellants: pressure rise is not, by itself, a reliable proxy for effective chamber energy conversion.

This work aims at acquiring deeper knowledge about solar system components including valuable materials as well as clearing viable ways towards extra solar systems. To achieve as much as possible the proposed objectives a viable Self-Sustaining Power Machine (SSPM), including an energy supply and propulsion paradigm shift is proposed. This system will replace the Nuclear Thermal Power and Nuclear Electric Propulsion Systems avoiding its inherent drawbacks, constraints and limitations. Therefore, this research addresses the need to firstly explore accessible celestial bodies within our solar system—planets, their moons and valuable accessible asteroids to utilize their resources. A primary challenge in this endeavor is ensuring a reliable energy supply. Consequently, this article proposes a disruptive change in both: energy supply and propulsion technology, replacing nuclear-based systems with SSPMs. These autonomous space vehicles would integrate the necessary SSPM-based resources to supply electrical power to all critical services and/or systems. Research results indicate us that the propellant fluid is crucial for achieving optimal results (high specific heat and low density with a high adiabatic expansion coefficient) ─(H₂). High temperature allows for high propellant expulsion velocity, resulting in high specific impulse and high thrust under lower propellant flow rate. Hydrogen (H₂) dramatically outperforms Nitrogen (N₂). At 3000K, H₂ achieves an exhaust velocity of 10.34 km/s (Isp ~1054 s), while N₂ only reaches 2.74 km/s (Isp ~280 s). H₂ produces significantly more thrust than N₂ under the same conditions. For example, at a mass flow rate of 100 kg/s and 3000K, H₂ generates ~103.4 tons of thrust, while N₂ generates only ~27.4 tons (Tables 17 & 20). With 50 tons of propellant to push a 10-ton spacecraft, H₂ heated to 3000K can achieve a final velocity of ~26.5 km/s. N₂ under the same conditions can only reach ~7.0 km/s, a massive difference in transit time and mission reach

Solid rocket motors (SRMs) have limited propulsion performance due to their short fuel combustion duration and low specific impulse. The supplementary combustion nozzle (SCN) injects air into the nozzle expansion section to promote secondary combustion of unburned exhaust gases, providing a novel solution to the aforementioned problems. This study fills the knowledge gap in coupled inlet-nozzle interactions using a comprehensive computational framework. The framework constructs the coupled internal and external flow fields of the SCN. It solves these fields using 2D axisymmetric compressible Navier-Stokes equations and the k-ε turbulence model. This allows for a systematic study of key inlet parameters, including outflow angle, outflow position, and mass flow rate. The obtained results demonstrate that an outflow angle of 0° provides a balance between the combustion efficiency and inlet total pressure recovery coefficient, representing the optimal configuration. A proximal inlet positioning near the throat increases the thrust gain, which results in inducing significant backpressure surges. This requires comprehensive design trade-offs taking into consideration that a peak net thrust gain occurs at an air-to-gas mass flow ratio of 1.23. After conducting a multi-parameter coupled optimization, the SCN achieves a net thrust gain of 8.18%, significantly outperforming conventional nozzles. This study provides theoretical insights for high-specific-impulse design in SRMs.

The initiation and stabilization of oblique detonation waves (ODWs) are critical to the practical implementation of oblique detonation engines (ODEs). In this study, the role of an on-wedge bump in facilitating ODW initiation within an ODE combustor is numerically investigated. The study employs two-dimensional reactive Euler equations with a simplified two-step chemical kinetic model to capture the flowfields and combustion characteristics of a fuel-rich hydrogen–air mixture at a flight altitude of 45 km. Without the bump, the ODE exhibits distinct initiation behaviors depending on the flight Mach number (Ma0). At low flight Mach numbers, the bump triggers wall-adjacent combustion or shock-induced combustion, while at high flight Mach numbers, it promotes a quasi-steady ODW, enhancing the maximum specific impulse (Isp) by up to 11.2% compared to the bump-free case. Furthermore, the influence of the bump pattern and geometric characteristics on wave structures and engine performance is assessed by introducing a nondimensional detachment intensity parameter, Ω, providing insights into forced initiation mechanisms for hypersonic propulsion. By demonstrating how on-wedge bumps enhance the combustion efficiency and stabilize detonation waves across varying Mach numbers, this work bridges a key gap between the idealized open-space studies and realistic confined combustor conditions.

An SPT-100-type electric thruster is simulated at nominal operating conditions using a two-dimensional, axisymmetric, fully kinetic particle-in-cell/Monte Carlo collision (PIC/MCC) plasma model. The model reproduces PIVOINE facility measurements, predicting a thrust and discharge current of 77.2 mN and 4.30 A versus 80.0 mN and 4.15 A measured, with thrust efficiency and specific impulse of 42.4% and 1575 s versus 45.0% and 1490 s. The power spectral density of the discharge current waveform reveals a dominant breathing-mode oscillation peak at 19.5 kHz and a secondary feature near 107 kHz. Time-resolved ion flux to the inner and outer radial walls exhibits quasi-periodic modulation and intermittent near-exit enhancements, demonstrating that instantaneous erosion drivers vary substantially over the oscillation cycle. Cross-correlation of the breathing-band (15–25 kHz) components indicates that the wall ion flux lags the discharge current by approximately 63° (≈ 9 µs, ≈ 18% of a breathing period), suggesting a delayed near-wall response associated with motion and reshaping of the ionisation/acceleration region and the near-exit sheath/field structure rather than ballistic ion time-of-flight. These results provide a phase-resolved pathway for connecting breathing-mode oscillations to wall ion bombardment in SPT-class Hall-effect thrusters.

Abstract Contamination control in inkjet printing cleanrooms is critical yet challenging due to the transient airflow disturbances caused by moving equipment. This study examines how exhaust geometry and suction momentum govern flow topology and particle clearance in such dynamic environments. A scalable computational framework is employed, combining a sharp-interface Immersed Boundary Method (IBM) on a block-structured Cartesian grid with Implicit Large-Eddy Simulation (ILES) and a two-way coupled PSI-cell particle model. Seven ventilation configurations are simulated, independently varying exhaust aperture, exhaust velocity, and supply-jet speed. Performance is quantified by end-of-cycle room balances (retention, removal, and escape ratios) and by a substrate-attached “critical zone” metric. Phase-resolved analysis links particle transport to specific flow structures, including capture layers, recirculation cells, and bypass paths. Results demonstrate that exhaust-side design is dominant: enlarging the outlet or increasing suction establishes a coherent capture layer that eliminates bed-top recirculation, reducing outward particle escape to approximately 1% and reducing critical-zone residue to near zero. In contrast, increasing supply-jet velocity enhances impingement and lateral entrainment, raising outward escape to 16-17% without improving clearance. The study reveals the mechanisms controlling dilute, drag-dominated particles and provides a transferable, non-dimensional protocol for chamber-scale ventilation design.

Electrospray thrusters promise compact, high specific impulse propulsion for small spacecraft, yet ground characterization remains confounded by secondary species emission and incomplete diagnostics of neutral products. To address these limitations, we perform energy-resolved mixed quantum/classical ab initio molecular dynamics of neutral 1-ethyl-3-methylimidazolium tetrafluoroborate colliding with Au extractor surfaces with impact energies from 10 to 100 eV to resolve fragment species spectra, charge states, kinetic energy partitioning, and scattering geometry. The simulations reveal a three-stage sequence with impact energy: the low energy regime, 10–20 eV, which favors ionic dissociation; intermediate energy regime, between 30 and 40 eV, opens a neutralization window; and high energy regime, greater than 50 eV, drives covalent fragmentation into many light products with mixed charge states. Fractional energy distributions show a transition from few-body, energy-concentrated outcomes in the low energy regime to many-body, energy-dispersed outcomes in the high energy regime. Deflection angle distributions exhibit a strong mass-to-angle anti-correlation such that heavier fragments favor small deflection, while lighter fragments are observed across the full deflection range but dominate larger deflection angles. The fraction of metastables peaks near 50 eV, coinciding with abundant neutral fragment production. Importantly, neutral bombardment still produces charged secondaries at the target even when the upstream ion plume is fully suppressed by a decelerating electrode. These findings provide a basis for de-biasing facility measurements by pairing tandem time-of-flight secondary ion mass spectrometry and a residual gas analyzer with suppression-bias corrections to inform the design of electrospray thrusters that reduce interception on extractor surfaces.

This work establishes a materials design paradigm that achieves simultaneous enhanced stability and performance for photon-based propulsion. We introduce an atomic-level bimetallic platform to overcome the inherent trade-off between high thrust efficiency and environmental stability, particularly against hydrolysis. This is achieved through bimetallic FeCu-MOFs synthesized via a one-step laser synthesis, where Fe3+ and Cu2+ co-crystallize with tricarboxylate ligands to form an isomorphous HKUST-1 derivative. This approach exploitshard-soft acid-base principlesto achieve several fundamental advances: enhanced bond strength, hydrolytic and thermal stabilitythrough the formation of robust Fe-O bonds, increasing water resistance by 20 times while preserving crystalline integrity; synergistic, delocalized energy dissipationvia d-orbital charge transfer (Fe3+→Cu2+), boosting uniform photothermal conversion to 91%; and inherent stoichiometric tunability, where the Fe:Cu ratio serves as a precise performance lever, providing a design strategy to optimize stability and performance. The optimized FeCu-MOF-M variant achieves record propulsion metrics-impulse coupling coefficient (191.80 µN/W), specific impulse (631.19 s), thrust density (61.86 µN/µg), and ablation efficiency (59.32%) -surpassing monometallic HKUST-1 by 15.7% and physical mixtures by 125%. By unifying hydrolysis resistance, efficient photothermal conversion, and atomic-level tunability, this stoichiometry-driven photothermal synergy bimetallic frameworks provides a solid foundation for next-generation energetic materials in demanding environments.

ABSTRACT Ammonium perchlorate (AP)-based solid propellants are widely used in booster rockets; however, AP combustion emits poisonous gases, including chlorine, which poses an environmental concern. To address these problems, oxidizers that are environmentally benign such as ammonium nitrate (AN) are applicable, but their application is restricted by their hygroscopic nature and incomplete combustion. A dual oxidizer system can solve these disadvantages. This research assesses the performance properties of the HTPB-based composite propellants prepared with various AN/AP ratios and catalytic additives, including ferrocene and vanadium pentoxide. Recipes that included and excluded additives were studied on burn rate, density, calorific value, specific impulse (Isp), mechanical properties, flame temperature, thermal decomposition behavior and hydrogen chloride (HCl) emissions. The dual oxidizer-ratio samples of 1.5:1 with ferrocene and vanadium pentoxide had 205.70% and 77.98% increase in the burn-rate and calorific-value, respectively, compared to the base formulation. Thermal analysis revealed that the 1.5:1 composition exhibited higher initial endothermic and slightly enhanced exothermic transitions as compared to the 1:1 ratio. Ferrocene and vanadium pentoxide reduced the chloride emission by 21.10% and 15.98%, respectively, in the experiment. Mechanical analysis also indicated that the addition of additives and the ratio of AN:AP had no significant effect on the strength, but acceptable flexibility and integrity could be achieved within standard mechanical limits of composite-propellant composites. Ferrocene was selected as the best additive, which provides significant combustion improvement and causes a decrease in the emissions harmful to the environment. These findings emphasize the gainfulness of streamlining both the oxidizer ratio and additive selection to the creation of the next generation, eco-friendly solid propellants to develop a superior propulsion system in space.

Ionic cocrystals (ICCs) hold significant potential to enhance performances for inorganic salts in fields of pharmaceuticals, photoelectricity, energetic materials, etc. However, the targeted design of ICCs with specific properties remains challenging. Herein, this work addresses the critical hygroscopicity of ammonium dinitramide (ADN), a vital green replacement for ammonium perchlorate (AP), by establishing the cooperative heterosynthon strategy. Specifically, improved binding energy in the cluster analysis revealed that moisture sensitivity originates from synergistic cation-anion interactions, which guided the identification of 3,4-diaminofurazan (DAF) as the optimal coformer through subsequent experimental screening. The resulting ADN/DAF cocrystal achieves unprecedented moisture resistance (critical relative humidity = 84.0% at 25 °C) among all ADN cocrystals, enables substitution for AP even surpassing ADN in specific impulse (272.25 s), and significantly enhances safety performance. Notably, this cocrystal demonstrates superior manufacturability: utilizing commercially available coformers and synthesizing via continuous flow production at the hectogram-scale with >88% yield. Therefore, this work presents an effective strategy for engineering moisture-resistant ionic cocrystals as well as showcasing high potential in applications for green energetic propellants.

ABSTRACT This study investigates the influence of Hydroxyl-Terminated Polybutadiene (HTPB) molecular weight (Mn), hydroxyl value, polydispersity index, and functionality on the mechanical and combustion properties of composite solid propellants (CSPs). Five samples (CSP-A to E) were prepared using HTPB grades with Mn ranging from 2462 to 4226 Da at a constant 68 vol% solid loading. Results revealed that moderate-molecular-weight HTPB formulations demonstrate optimal overall performance. Specifically, CSP-C (Mn = 3253 Da) exhibited the highest energetic output, with increases of approximately 5–7% in specific impulse (271.3 s), 7–10% in density impulse (458.5 s), and ~15% in heat of decomposition (7.08 MJ/kg), respectively, compared to other samples. While lower molecular weight and higher functionality enhanced tensile strength (up to 6.81 kgf/cm²) through increased crosslink density, excessive crosslinking led to reduced ductility. These findings highlight the importance of optimizing intrinsic HTPB characteristics to tailor CSP performance.