Specific Impulse Isp Research Articles

Synergetic and Performance Characteristics of A High-Speed Pre-Cooled Propulsion Concept

Abstract Pre-cooled air-breathing cycles are promising candidates to power future high-speed flight as well as Single-Stage-To-Orbit vehicles, due to their increased efficiency over contemporary propulsion systems and launch vehicles. These concepts usually feature complex interactions in the synergy of their thermodynamic cycles. In this study, a performance model of such a cycle is developed for its air-breathing mode of operation. One-dimensional thermodynamic modeling is employed within a component-level approach, to evaluate the performance and operation of the cycle under investigation in the range of 1.35 = ≤ 8 = 5 and conditions of up to 26 kilometers altitude. The model is validated quantitatively and qualitatively for both design and off-design conditions. The specific impulse Isp and specific thrust, as predicted by the model, agree within less than 5% for both design and off-design point conditions, while it captures the trend of Isp for the range modeled. Moreover the maximum gross thrust point is predicted correctly at M∞ = 4. The fundamental operating principles and synergetic characteristics of the engine at design and off-design conditions are investigated and reported. A model which does not feature a bypass duct is created and compared for the same inflow conditions and mission profile. It is found that the engine without the bypass duct exhibits reduced specific impulse up to 32% lower at off-design conditions while the overall trend of engine efficiency cannot be properly captured without modeling of the bypass duct, especially at the region of M∞ < 3.5.

Development of hydrogen peroxide based micro thruster utilizing biomass-based catalyst bed

Micro thrusters, also referred to as micro-newton thrusters, are thrusters having a wide range of force capabilities in the micro-newton range. Micro thrusters are used in CubeSats, tiny robots and drones, micro- and nanosatellites, spacecraft attitude control, etc. The aim of the present study is to develop a monopropellant thruster (MPT) for 90 % concentrated hydrogen peroxide (H2O2). A rotatory evaporator was used to obtain a 90 % H2O2 form the available concentration of 50 % H2O2. The next stage involved the designing of the monopropellant thruster utilizing the NASA CEA rocket performance algorithm and the derived mathematical model. The test facility was constructed which includes a thruster, a feed line system, a static test firing platform, a data collecting system and the measuring sensors. To measure the thrust of the thruster, a pendulum thrust mechanism was constructed and fabricated into an experimental test stand. Lastly, the MPT chamber, which uses stainless steel mesh, was prepared and packed with the biomass-based catalyst. The 14 bar injection pressure conditions were used for the fire tests. With a mass flow rate of 5 g/s and a decomposition pressure of 10 bar, the theoretical specific impulse (Isp) was calculated to be 1993.2 Ns/kg at an O/F ratio of 0.002. The test firing with 90 % concentrated hydrogen peroxide gave an experimental specific impulse of 794.5 Ns/kg, which gives efficiency of 39.9 %.

Tetrazole-grafted-hydroxyl terminated polybutadiene: A novel energetic binder for solid rocket propulsion

The present study reports on the synthesis, characterisation, and energetic aspects of poly-nitrogen containing azole-grafted hydroxyl terminated polybutadiene (HTPB) based energetic binders (EBs). A comprehensive computational analysis at the B3LYP/6-311++G(d,p) level of theory reveals that azole-grafted HTPB variants exhibit significantly higher HoF (1061.6 to 1659.8 kJ/kg) and densities (1.26 to 1.61 g/cc) as compared to traditional non-energetic binders such as HTPB (specific heat of formation (HoF): 781.8 kJ/kg, density: 0.92 g/cc),. Their ballistic properties, including specific impulse (Isp: 266.2 to 283.8 s) and density specific impulse (ρIsp: 484.5 to 547.4 g/cc.s) are also superior as compared to HTPB (Isp: 283.8 s, ρIsp: 547.4 g/cc.s). Moreover, the synthetic methodology for grafting azole compounds onto HTPB, utilizing cost-effective, metal-free, and environmentally sustainable procedures with straightforward iodine and base mediation, underscores its potential for large-scale industrial applications. In addition to its advantages in synthesis and energetic properties, tetrazole-grafted HTPB exhibits favourable physicochemical properties: it has a low molecular weight ([Mn(VPO)] = 3023), high specific gravity (1.25 g/cc at 30 °C), moderate flowability (η = 623.6 Pa.s at 30 °C and 62.7 at 60 °C), a low glass transition temperature (−57 °C), mechanical insensitivity (low impact sensitivity and no ignition up to 400 °C), high thermal stability (onset temperature, Td,onset = 190 °C, TDSC = 246 °C), and a desirable hydroxyl value (47 mg/KOH). These properties make tetrazole-grafted HTPB an appealing choice for advancing novel propellant composites. Overall, these findings not only pave the way for advancements in HTPB-based propellants but also inspire further exploration in real-world applications, highlighting the transformative potential of azole-grafted HTPB in propellant technology.

Numerical Study of Porous Cylindrical Containment for the Fuel Element of a Centrifugal Nuclear Thermal Rocket

For deep-space propulsion and interplanetary exploration, the centrifugal nuclear thermal rocket (CNTR) has the ability to achieve a very high specific impulse (Isp) metric beyond that of conventional chemical rockets or solid-core nuclear thermal propulsion systems. The high Isp allows the rocket to use less propellant or achieve a higher velocity for shorter transit times. However, the cylindrical containment structure of a CNTR fuel element encounters extreme conditions, as it houses molten uranium at temperatures exceeding 1408 K, leading to challenges such as dissolution, chemical reactions, and thermal stresses that conventional materials struggle to withstand. This study aims to address this issue by analyzing appropriate materials for constructing the cylindrical containment component. The operating conditions of the annular porous medium that confines the liquid uranium in the centrifugal fuel element are simulated by conducting a comprehensive one-dimensional numerical analysis using a range of candidate porous materials, including Mo, W, zirconium carbide, and silicon carbide. The porous structure facilitates the flow of the hydrogen propellant into the internal molten uranium section, where it gains significant thermal energy while simultaneously cooling the cylinder. The containment cylinder has an internal temperature of 1478.1 K, exceeding the melting point of uranium, while the external gas temperature of the hydrogen propellant is much lower. This temperature difference induces significant thermal stresses in the cylinder. The porous containment cylinder made from molybdenum was able to maintain elastic deformation throughout the thickness of the cylinder, showcasing its ability to handle these extreme thermal stress conditions. Tungsten, on the other hand, experienced plastic deformation at the cylinder’s edges and elastic deformation through the middle radial locations. In contrast, the stresses experienced by the ceramic materials far exceeded their failure stress values, leading to brittle failure. These findings will help with the refinement of the CNTR design, edging it closer to practical implementation.

Construction and Modification of Nitrogen-Rich Polycyclic Frameworks: A Promising Fused Tricyclic Host-Guest Energetic Material with Heat Resistance, High Energy, and Low Sensitivity.

The construction and modification of novel energetic frameworks to achieve an ideal balance between high energy density and good stability are a continuous pursuit for researchers. In this work, a fused [5,6,5]-tricyclic framework was utilized as the energetic host to encapsulate the oxidant molecules for the first time. A series of new pyridazine-based [5,6] and [5,6,5] fused polycyclic nitrogen-rich skeletons and their derivatives were designed and synthesized. Two strategies, amino oxidation and host-guest inclusion, were used to modify the skeleton in only one step. All compounds exhibit good comprehensive properties (Td(onset) > 200 °C, ρ > 1.85 g cm-3, Dv > 8400 m s-1, IS > 20 J, FS > 360 N). Benefiting from the pyridazine-based fused tricyclic structure with more hydrogen bonding units and larger conjugated systems, the first example of [5,6,5]-tricyclic host-guest energetic material triamino-9H-pyrazolo[3,4-d][1,2,4]triazolo[4,3-b]pyridazine-diperchloric acid (10), shows high decomposition temperature (Td(onset) = 336 °C), high density and heats of formation (ρ = 1.94 g cm-3, ΔHf = 733.4 kJ mol-1), high detonation performance (Dv = 8820 m s-1, P = 36.2 GPa), high specific impulse (Isp = 269 s), and low sensitivity (IS = 30 J, FS > 360 N). The comprehensive performance of 10 is superior to that of high-energy explosive RDX and heat-resistant explosives such as HNS and LLM-105. 10 has the potential to become a comprehensive advanced energetic material that simultaneously satisfies the requirements of high-energy and low-sensitivity explosives, heat-resistant explosives, and solid propellants. This work may give new insights into the construction and modification of a nitrogen-rich polycyclic framework and broaden the applications of fused polycyclic framework for the development of host-guest energetic materials.

Analysis of Film Cooling for a 3kN LOX/butane Demonstrator Engine

Regenerative cooling has been the primary cooling method for every modern launch vehicle engine, except for the Viking: a film-cooled (with ablative throat) N2O4 / UH25 engine used on the first stage of the Ariane rockets 2-4. Despite this, film-cooling as a stand-alone cooling method has traditionally been considered insufficient for the high combustion temperatures and long burn times associated with launcher engines. This study explored the feasibility of a solely film-cooled engine at the demonstrator scale (3 kN), as a prototype for a lightweight launcher engine. A wide range of liquid oxygen (LOX)/butane engines were modelled and from this a relationship was determined to predict chamber wall temperature for a given oxidiser-to-fuel ratio (O/F), chamber pressure, and amount of film cooling. Notably, this equation was found to apply to both a 3 kN and a 30 kN engine. Numerical modelling of engine specific impulse (Isp) using this equation then found the conditions yielding optimal engine performance.

Numerical Investigation on the Effect of Fuel-Rich Degree in the RBCC Engine under the Ejector Mode

The ejector mode of the Rocket-Based Combined-Cycle (RBCC) engine is characterized by high fuel consumption. This study is aimed at investigating the influence of the rocket fuel-rich degree on the RBCC engine’s performance under the ejector mode combined with simultaneous mixing and combustion (SMC). Numerical simulations were conducted for various rocket mixing ratios (Φ=1.6~3.2) under subsonic (Maf=0.9) and supersonic (Maf=1.8) flight conditions. It was observed that a high fuel-rich degree in the rocket plume negatively impacts the eject performance under all conditions. However, it improves the overall performance (Isp) at high flight Mach numbers (Maf). For supersonic conditions, increasing the fuel-rich degree promotes greater fuel participation in combustion, thereby enhancing RBCC engine performance. Nevertheless, the subsonic-supersonic mixing layer exhibits low evolution, resulting in a decrease in reaction efficiency from 29.2% to 12.0% as the Φ decreases from 3.2 to 1.6. Consequently, there is an inefficient utilization of fuel. To optimize RBCC engine performance, the rocket fuel-rich degree can be appropriately increased. However, this increase should be limited to prevent fuel wastage arising from low reaction efficiency. Under subsonic conditions (Maf=0.9), the low kinetic energy of captured air leads to the occurrence of “negative thrust surface” and “wall impact” phenomena, which hinder the efficient and stable operation of the RBCC engine. Consequently, adjusting the fuel-rich degree alone cannot promote specific impulse (Isp), and a low fuel-rich degree is considered an ideal strategy when combined with adjustable nozzle technology.

A Solar Thermal Steam Propulsion System Using Disassociated Steam for Interplanetary Exploration

Sustainable space exploration will require using off-world resources for propellant generation. Using off-world-generated propellants significantly increases future missions’ range and payload capacity. Near Earth Objects (NEOs) contain a range of available resources, most notably water-ice and hydrated minerals. However, water-bearing regolith needs to be excavated and the water extracted. Water is a compelling choice for fuel as it is readily available in interplanetary space and easily stored. In this paper, we propose using solar concentrators, which can efficiently convert incident sunlight into heat without the need for moving parts. When water is heated up to 4000 K, a value consistent with high-performance refractive materials, it experiences significant disassociation into H2, O2, OH, H, and O components, providing a path for adding considerable additional chemical energy per degree of temperature increase, and producing theoretical specific impulse (Isp) values in the range of 643 s to 659 s.

2,2'-Azobis(1,5'-bitetrazole) with a N10 Chain and 1,5'-Bitetrazolate-2N-oxides: Construction of Highly Energetic Nitrogen-Rich Materials Based on C-N-Linked Tetrazoles.

A series of high-nitrogen compounds, including a unique molecule 2,2'-azobis(1,5'-bitetrazole) with a branched N10 chain and 1,5'-bitetrazolate-2N-oxides, were synthesized successfully based on C-N-linked 1,5'-bistetrazoles using azo coupling of N-amine bonds and N-oxide introduction strategies. All compounds were characterized by NMR spectroscopy, IR spectroscopy, elemental analysis, and differential scanning calorimetry, in which the structures of five compounds were further determined by single-crystal X-ray diffraction analysis (2, T-N10B, 3a, 3b, and THX). The nitrogen contents of these five compounds range from 63.62 (THX) to 83.43% (T-N10B), which are much higher than that of CL-20 (38.34%). The heat of formation for the prepared compounds was calculated by using the Gaussian 09 program, with T-N10B having the highest value of 5.13 kJ g-1, about 6 times higher than that of CL-20 (0.83 kJ g-1). The calculated detonation performances by EXPLO5 v6.05.04 show that THX has excellent detonation performance (D = 9581 m s-1, P = 35.93 GPa) and a remarkable specific impulse (Isp = 284.9 s).

Effect of plasma confinement on laser ablation propulsion parameters by using external semi-elliptical cavities for Aluminum and Silver propellants

Ablative laser propulsion (ALP) is a type of beam-powered propulsion in which thrust is produced by ablating a target using a high-intensity laser beam. In this research work, we calculated the ALP parameters momentum coupling coefficient (Cm) and specific impulse (Isp) by confining the plasma with the help of external semi-elliptical cavities. The cavity minor axis was fixed at 2.5 mm while the major axis varied to 7 mm, 9 mm, 11 mm, 13 mm and 15 mm. Silver and aluminum samples were used as propellants. A pulsed Q-smart Nd:YAG laser operating at 532 nm wavelength was used to ablate the target. A significant increase in the value of Cm and Isp was observed using confined ablative propulsion. The highest Cm and Isp were recorded for the 9 mm major axis cavity, for which Cm was increased from 5.01 × 10−5 N-s/J to1.75 × 10−4 nearly 249% for silver and from 3.4 × 10−5 N-s/J to 8.9 × 10−5 N-s/J nearly 161% for aluminum. Hence the use of semielliptical cavities is an effective technique for enhancing propulsion parameters. Cavity acts as a trap for plasma and shockwaves. Confinement of plasma and reflection of shockwaves from the cavity walls results in enhanced laser ablation propulsion parameters.

Experimenting on semiconductors using ablative laser propulsion to investigate propulsion parameters

A laser can be used to propel distant objects. The ability of silicon to produce thrust and its propulsive parameters are needed to be studied in ablative laser propulsion (ALP). In this work, pure silicon and silicon doped with indium were subjected to ALP to achieve the momentum coupling coefficient and specific impulse to investigate its worth as a propellant. The experiment was conducted using the Nd:YAG laser (Quantel Brilliant) operating at fundamental harmonic (λ = 1064 nm and 5 ns pulse duration). In the given range of fluence, 1 × 105–5 × 105 J/m2, no significant difference among both samples for values of momentum coupling coefficient (Cm) and specific impulse (Isp) is observed; for instance, both propellants follow a decreasing trend for Cm. However, maximum enhancement for Cm and Isp is observed with a cavity aspect ratio of one. Cm and Isp are enhanced about 1.75 and 4.5 times, respectively, for pure silicon. The external cavity does not have any impact on Cm values for indium-doped silicon while values of Isp showed considerable enhancement.

Advanced Electric Propulsion Concepts for Fast Missions to the Outer Solar System and Beyond

Electric Propulsion (EP) comprises all types of space propulsion in which a certain amount of propellant is ionized and then accelerated by electric or magnetic fields, or both. This propulsion technology allows for much higher specific impulses (Isp > 2,000 s) than conventional chemical propulsion (Isp < 500 s), resulting in a major reduction of the propellant mass or a considerably higher final speed for a certain space mission. Hence, EP can enable very challenging space missions as DAWN has clearly shown. Furthermore, an EP system coupled with an advanced nuclear reactor could enable fast manned missions to Mars (one-way travel times less than 4 months). This propulsion technology can be scaled up to even higher specific impulses (Isp > 5,000s). However, the power needed for the same thrust is also increasing. An Oberth maneuver performed very close to the Sun could provide the additional power to a high-Isp Solar Electric Propulsion (SEP) system in order to reach the needed delta-v for challenging interstellar precursor missions. A breakthrough in power source specific mass is needed in order to enable missions with ultra-high specific impulses (> 10,000 s); this breakthrough could be realized having the power source not on board, as with Laser-powered Electric Propulsion (LEP), where the needed power is beamed to the spacecraft from an external laser source. In this case the on-board power source is limited to a light-weight photovoltaic receiver/converter. The development of ultra-high Isp ion thrusters powered by an external laser source could enable the most challenging interstellar precursor missions up to the Oort Cloud and beyond. This paper gives an update on the status of these advanced propulsion concepts, and provides some examples of interstellar precursor missions enabled by advanced EP systems which could be launched before 2040. Keywords: Interstellar, Electric Propulsion, Oberth, Laser

Transmissive Mode Laser Micro-Ablation Performance of Ammonium Dinitramide-Based Liquid Propellant for Laser Micro-Thruster.

The transmissive mode laser micro-ablation performance of near-infrared (NIR) dye-optimized ammonium dinitramide (ADN)-based liquid propellant was investigated in laser plasma propulsion using a pulse YAG laser with 5 ns pulse width and 1064 nm wavelength. Miniature fiber optic near-infrared spectrometer, differential scanning calorimeter (DSC) and high-speed camera were used to study laser energy deposition, thermal analysis of ADN-based liquid propellants and the flow field evolution process, respectively. Experimental results indicate that two important factors, laser energy deposition efficiency and heat release from energetic liquid propellants, obviously affect the ablation performance. The results showed that the best ablation effect of 0.4 mL ADN solution dissolved in 0.6 mL dye solution (40%-AAD) liquid propellant was obtained with the ADN liquid propellant content increasing in the combustion chamber. Furthermore, adding 2% ammonium perchlorate (AP) solid powder gave rise to variations in the ablation volume and energetic properties of propellants, which enhanced the propellant enthalpy variable and burn rate. Based on the AP optimized laser ablation, the optimal single-pulse impulse (I)~9.8 μN·s, specific impulse (Isp)~234.9 s, impulse coupling coefficient (Cm)~62.43 dyne/W and energy factor (η)~71.2% were obtained in 200 µm scale combustion chamber. This work would enable further improvements in the small volume and high integration of liquid propellant laser micro-thruster.

Numerical analysis of LOx-BioLPG combustion in high-pressure liquid rocket engine propulsion system

The present work numerically investigates the steady state LOx-BioLPG combustion in a high-pressure liquid rocket engine propulsion system, based on the RANS approach using Eulerian single-phase thermo-chemical modeling with Peng-Robinson equation of state. The study primarily focuses on the less carbon emission and the economical range of operation for various equivalence ratios. The critical equivalence ratio (Φ) is predicted to be 0.62, below which the combustion does not initiate. The maximum CO emission in combustion is found at Φ = 2.32, and the maximum specific impulse (Isp) is obtained 327.92 at Φ = 1.50. However, for clean burning of the fuel and less carbon footprint, lean mixture combustion is desirable. The CO emission for Φ = 1.50 is 41%, which is considerably a higher carbon emission percentage for this combustion process. On the other hand, the lean mixture at Φ = 0.80 has a CO emission of 14.7% at the nozzle outlet which is relatively small. The developed power is 57.73 MW with an Isp of 310.96 at Φ = 0.80 which is close to the Isp obtained for Φ = 1.50. Thus, for eco-friendly clean burring, combustion of LOx-BioLPG fuel for Φ = 0.80 is found to be the best in this analysis.

Feasibility analysis of solar thermal propulsion system with thermal energy storage

Possessing relatively high specific impulse and moderate thrust levels, solar thermal propulsion (STP) is a promising candidate in spacecraft propulsion system. However, the traditional solar thermal propulsion system suffers from thrust failure in the shadow area, which seriously affects its applicability. In this paper, we investigate feasibility of regenerative solar thermal propulsion system (RSTP) incorporating thermal energy storage, which can effectively overcome unmatched synchronous working time and illumination time. A numerical model for RSTP considering the whole energy transfer process from light concentrating, heat storage, to thrust generation is built, which is verified by experiment measurements with relative errors less than 15 %. The result shows that the maximum time to complete heat storage is about 4000 s, which is within the illumination time for low Earth orbit. In the solar eclipse region, the thrust (Ft) and the specific impulse (Isp) of the system increase with the propellant flow rate, which can reach about 2 N and 690 s, respectively. What’s more, the system can operate for around 100 s continuously at the maximum thrust in the shadow area. This work provides alternative approaches for microsatellite propulsion with high specific impulse, high thrust, and continuous operation despite presence of solar eclipse.

Thermolysis and sensitivities of solid propellants using characterized nano oxidizers involving energy performance evaluation

This investigation was devoted to exploring the thermal decomposition and sensitivity characteristics of solid propellants using nano-AN and nano-AP as oxidizers. The morphology, surface elements, and crystal structure of nano-AP and nano-AN were investigated by SEM, BET, EDS, and XRD. DSC and TG-MS were used to investigate the thermal decomposition mechanism of nano-AP and nano-AN solid propellants. The mechanical sensitivity of the different solid propellants containing nano-AN and nano-AP was tested and their energy performance was also evaluated. The results show that the microscopic morphology of nano-AN and nano-AP is network-like with the one-dimensional scale less than 100 nm, and the crystal phases of the two are consistent with the raw AN and raw AP, respectively. The thermal decomposition activation energy of AN/CMDB-4 propellant containing nano-AN is 717.46 kJ/mol, while that of AP/HTPB-6 propellant containing nano-AP is 422.33 kJ/mol. The main decomposition products of AN/CMDB-4 propellant are CO2, N2O, NO, CH2O, CO, N2, H2O, CH4 and H2. The products of CO2, N2O, NO, CH2O, CO, N2, H2O, CH4 and H2 are also generated during the decomposition of AP/HTPB-6 propellant. With the increase of nano-AN and nano-AP in propellants, the mechanical sensitivity of AN/CMDB solid propellant decreases, while that of AP/HTPB solid propellant increases. The theoretical standard specific impulse Isp of AN/CMDB and AP/HTPB propellants are 2439.5 N·s/kg and 2449.3 N·s/kg, respectively. The above conclusions show that nano-AP and nano-AN can improve the thermal decomposition performance and the safety of solid propellants, which are expected to be widely used in solid propellants.

Demonstration of electric micropropulsion multimodality.

Electric propulsion has become popular nowadays owing to the trend of miniaturizing the size and mass of satellites. However, the main drawback of the most popular approach—Hall thrusters—is that their efficiency and thrust-to-power ratio (TPR) markedly deteriorate when its size and power level are reduced. Here, we demonstrate an alternative approach—a minute low-power (<50 W), lightweight (~100 g), two-stage propulsion system. The system is based on a micro-cathode vacuum arc thruster with magnetoplasmadynamic second stage (μCAT-MPD), which achieves the following parameters: a thrust of up to 1.7 mN at a TPR of 37 μN/W and an efficiency of ~50%. A μCAT-MPD system, in addition to “traditional” inverse, displays the anomalous direct (growing) “TPR versus specific impulse Isp” trend at high Isp values and allows multimodality at high efficiency.

High-temperature superconductor-based power and propulsion system architectures as enablers for high power missions

The increasing competitiveness of electric propulsion systems (EPS) for primary spacecraft propulsion has paved the way for higher payload mass fractions by offering significantly higher specific impulses than chemical systems. Concurrently, High-Temperature Superconductors (HTS) have reached an unprecedented level of industrial maturity in recent years, and considering their low masses, compactness, and high current densities, they offer the potential to act as a disruptive technology in several spaceflight applications such as power management systems, re-entry and radiation shielding as investigated in the EU-funded MEESST project, as well as EPS. In the latter case, efforts are already ongoing to develop an HTS-enhanced Applied-Field Magnetoplasmadynamic (AF-MPD) thrusters for high power mission applications. The Tsiolkovsky equation infers that the payload mass fraction increases indefinitely with increased Specific Impulse (Isp), however, in the case of electric propulsion, the dependence of thrust on the available power complicates this issue when transfer time is a primary driver. Here, the Tsiolkovsky equation becomes inadequate and considering a non-dimensional version of the Tsiolkovsky equation in terms of the mission ΔV and transfer time, the EPS thrust efficiency, and the specific mass of the power system becomes necessary. This paper first discusses the recent advances in HTS and their suitability for spaceflight, before reviewing The development of power system technologies is reviewed and a conceptual power system architecture incorporating HTS is presented. These technologies are analysed using a non-dimensional Tsiolkovsky approach and their impacts on the overall payload mass fraction are assessed. For high-power missions (>100 kW), the use of HTS is shown to have a highly beneficial impact on the mass of the power system. Correspondingly, this enables higher payload mass fractions achievable at increased specific impulse operation, thus strengthening the case for high-power, high-Isp EPS technologies such as AF-MPDT.

Interface Adhesion Property and Laser Ablation Performance of GAP-PET Double-Layer Tape with Plasma Treatment.

In the field of laser ablation micro-propulsion, the property of double-layer tape has significant impact on the propulsion performance. In this paper, low temperature plasma was used to treat the surface of polyethylene terephthalate (PET) to improve its adhesion with energetic polymer. The PET surface pre- and post-plasma treatment was characterized by X-ray photoelectron spectroscopy (XPS) and atomic force microscopy (AFM), and the enhancement mechanism of the interface adhesion was discussed. In addition, the ablation performance of the double-layer tape after the plasma treatment was studied. The results showed that the plasma etching effect increased the root mean square roughness of the PET surface from 1.74 nm to 19.10 nm. In addition, after the plasma treatment, the number of C–OH/COOH bonds and O=C–O bonds increased, which also greatly improved the adhesion between the PET and energetic polymers. In the optimization of the ablation performance, the optimal laser pulse width was about 200 μs. The optimal values of the specific impulse (Isp), impulse coupling coefficient (Cm), and ablation efficiency (η) were 390.65 s, 250.82 μN/W, and 48.01%, respectively. The optimization of the adhesion of the double-layer tape and the ablation performance lay the foundation for the engineering application of laser ablation micro-thrusters.

Salt Formation, to Realize a Good Combination of High Energy and Low Sensitivity of Nitroform-Based Energetic Compounds

Energetic salts constitute an important issue in the field of energetic materials. In this research, the series of energetic salts 5–8 with nitroform structures were obtained by a single-displacement reaction, and their crystal structures were fully characterized by X-ray single-crystal diffraction. As energetic compounds, their energy performance and sensitivity to mechanical stimuli have been explored. These compounds not only exhibit excellent detonation performance (Dv = 8.45–9.13 km s–1 and P = 30.5–37.4 GPa) and extremely high specific impulse (Isp = 262.63–277.42 s) but also show low sensitivity to mechanical stimuli. According to the crystallographic data, the relationship between molecular structure and sensitivity has been explored. The results show that the positive electrostatic potential near the nitroform and the interaction between the intermolecular nitro oxygen atoms (O···O interaction) are significantly reduced after salt formation. In addition, this trend of change is consistent with the sensitivity distribution of these energetic salts (the impact sensitivity decreased from 5 to 15 J and the friction sensitivity decreased from 100 to 240 N). This study not only provides ideas for the molecular design of high-specific-impulse energetic compounds but also reveals the relationship between the structure and sensitivity of nitroform energetic compounds from the perspective of crystallography.