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Compact micronewton thrusters can provide attitude control for small satellites to increase mission duration and enable constellation flying. Micronewton thrust control can also enhance missions that require precision pointing such as space telescopes, laser interferometers, and laser communication relays. Film-Evaporating MEMS Tunable Array (FEMTA) is a novel micropropulsion device that generates thrust by heating a micron-scale water capillary to induce controlled film-evaporation. Thrust stand tests under high vacuum have shown that FEMTA can produce controllable thrust of 150 μN at 70 s specific impulse using 0.65 W of electrical power and ultra-pure deionized water as propellant. An undesired quiescent propellant loss rate is inherent to the current FEMTA design which limits its life span and reliability. To derive mitigations to this issue, the behavior of the fluid interface within the FEMTA micronozzle was characterized through direct experimentation. A test-bed FEMTA design was created which enabled direct observation of the liquid within the micronozzles and precise control over critical nozzle dimensions. These test-bed devices were used to measure contact angle, Laplace pressure, and total quiescent propellant loss rates for multiple nozzle configurations. Finally, a next generation FEMTA design was derived from the findings of these studies and its propulsive performance was measured under high vacuum on a micronewton thrust stand. Microfabrication was performed at Purdue’s Birck Nanotechnology Center and vacuum testing was completed at Purdue’s High Vacuum Lab.

Energy released during antimatter annihilation can be used to heat a working fluid to directly produce thrust (an “antimatter rocket”) or run a thermodynamic cycle to generate power for conventional electric propulsion systems (antimatter power generation). For a solid reactor core, we evaluate and compare such antimatter propulsion concepts and assess the resulting performance for different propellants (hydrogen, water, methane, and carbon dioxide) and electric propulsion technologies (arcjets, Hall thrusters, and gridded ion thrusters). In contrast to positron–electron annihilation (which only produces gamma rays), complete stopping of relativistic pions released during antiproton–proton annihilation increases the required reactor core size by approximately an order of magnitude, resulting in significantly larger propulsion systems. Because of reactor core material considerations, the maximum temperature and specific impulse achievable in antimatter rockets is limited. By instead operating the core at a lower temperature to drive a closed thermodynamic Brayton cycle, antimatter power generation can run electric propulsion systems with higher specific impulses. Although this reduces the propellant mass needed for a given mission, the combination of thermodynamic cycle efficiency (around 30%) and electric thruster efficiency (typically between 30%–70%) increases the antimatter mass required by one to two orders of magnitude. The lower thrust of electric propulsion systems also leads to longer burn and maneuver times. Considering emerging industry trends, the use of renewable propellants offers large potential benefits to future antimatter propulsion systems that can refuel in space.

This paper investigates whether biased electrodes can improve the performance of a microwave-discharge water ion thruster, which has been proposed as a propulsion system for CubeSats. Previously, for several microwave ion sources, the beam currents have been increased by mounting a biased electrode in the discharge chamber. The effect of a biased electrode on a water ion thruster was unclear because its magnetic field profile and characteristic length compared with the microwave wavelength are different from those used in previous studies. To investigate the effect, a thruster was divided into three parts: magnets, a microwave antenna, and a discharge chamber, which is comprised of an electrically conductive sidewall and a screen grid. The parts were isolated electrically. Bias voltages from − 44 to ＋ 44 V were applied to the microwave antenna or magnets by an additional power supply. These bias voltages were applied with respect to the screen grid and sidewall voltage. A maximum screen current of 12.7 mA was obtained for an antenna bias voltage of 38.2 V. This is 2.3 times larger than the current without a bias voltage. This increase could be caused by a combination of an increase of plasma density and ion extraction efficiency. The plasma density could increase with the gain of the total absorbed power. The ion extraction efficiency could increase due to the change in the plasma potential distribution. Overall, the mass utilization efficiency of the water ion thruster increased from 3.3% to 7.9% and the discharge loss decreased from 280 to 150 W/A at the antenna bias of 38 V. The performance with water electron source was evaluated. The thrust force, specific impulse, thrust-to-power ratio, thrust efficiency were 190 μ N, 350 s, 5.9 μ N/W, and 1.0%, respectively at the antenna bias of 25 V.

High energetic nanoparticle as additives is of great significance for liquid hydrocarbon fuels that will exhibit high heat value and potential performance improvement in hypersonic and related combined circle propulsion system. In the study, the JP-10 slurry fuel containing 16 wt% aluminum nanoparticles was experimented upon a scramjet engine to investigate the combustion efficiency and engine performance in comparison to pure JP-10 fuel. It is the first to realize the stable combustion of hydrocarbon fuel containing nanoparticles in scramjet engine. The wall pressure profile and calculated Mach number along the combustor are firstly compared and analyzed with the equivalence ratio of 0.56, 0.73 and 0.91. At the equivalence ratio of 0.56, the addition of aluminum nanoparticles promoted the engine to transition from the scramjet mode to the dual-model ramjet mode in advance, and the wall pressure along the combustor path to increase significantly. The combustion efficiency data of the two fuels are calculated by one-dimensional impulse analysis method. The results show that the combustion efficiency of slurry fuel is significantly higher than that of pure fuel. Besides, the addition of aluminum nanoparticles also enhanced the density specific impulse. When the equivalence ratio is 0.56, 0.53 and 0.91, the density specific impulse of slurry fuel increases by 34.37%, 11.33% and 8.96% respectively. The above research indicates that the use of slurry fuel containing aluminum nanoparticles indeed improves engine performance and has great potential in engineering applications.

Using a permanent magnet to realize a design with a large height-radius ratio can effectively reduce the radial size of the thruster. The experimental results show that a Hall thruster with a large height-radius ratio also has significant advantages in its discharge performance. Taking the 1.35 kW Hall thruster HEP-1350PM, which has a large height-radius ratio, as an example, the thrust, specific impulse, and efficiency of HEP-1350PM are increased by up to 4.3%, 10.4%, and 7%, respectively, as compared to SPT-100, while its weight is only 56% of that of SPT-100. Numerical simulations were performed to understand the discharge characteristics of a Hall thruster with a large height-radius ratio and the internal mechanism of its performance improvement. The results show that the small middle diameter and wide channel structure concentrate the distribution of the neutral gas in the channel, accelerate the decrease of the propellant density along the axial direction, and move the main ionization zone closer to the channel upstream. These factors increase the ionization concentration, ionization rate peak, working medium utilization rate, and voltage utilization rate. They also reduce the ion loss on the wall surface and plume divergence, thereby improving the comprehensive discharge performance. This study verifies the feasibility of a large height-radius ratio in the design of a Hall thruster, and demonstrates that the design would not only be valuable for reducing the volume and weight of the thruster, but also has the advantages of enhancing its discharge performance and prolonging its working life.

This study presents a broad comparison of samples, which are made of polytetrafluoroethylene (PTFE) filled with different dopants according to the same preparation steps, as propellants on a laser-electric hybrid thruster with a common design base. The impulse bit, specific impulse, efficiency and discharge success rate of the propellant samples were measured and analyzed under the same operation. The experimental results illustrated that the dopants of alloy could improve the impulse bit, that of salt could promote the stability of the discharge process, and the samples doped with oxide performed dramatic specific impulse and thrust efficiency. And three representative samples were adopted to investigate the dependence of propulsion characteristics on different working parameters in the respective doping ratios. It is found that the sample doped with TiO2 in lower doping rate was expected to rank maximal specific impulse and efficiency at most operations, especially under the conditions of low laser energy, long discharge width and high initial voltage. And the samples doped with Co–Cr alloy in low doping ratio appeared better performance at low discharge width and laser energy, while those in high doping ratio performed greater performance at high discharge width and laser energy.

This paper developed and tested a micro air-fed magnetoplasmadynamic thruster for small satellites. The thruster is a type of miniaturized electric propulsion system designed base on conventional full scale magnetoplasmadynamic thruster that operates at hundreds of kilowatts power. The thruster is designed and tested using normal air as the propellant under the pulse operation mode on a calibrated micro-force measurement thruster stand. The experiments revealed that the thruster can generate a 34.534 μ Ns impulse bit with an average power input of 1.857 ± 0.0679 W and thrust to power ratio of 8.266 μ N/W, and the specific impulse is calculated to be 2319 s with the thruster efficiency of 9.402%, which is quite competitive comparing with other solid-state and liquid-fed pulse-mode thrusters. This paper presents the design and test results for the thruster under a low power level as well as an analysis of its problems and limitations with corresponding future research and optimization directions noted at the end.

This paper presents the design methodology and performance testing of an additively manufactured resistojet operating on hydrogen as propellant. Additive manufacturing allows to produce complex monolithic resistors, resulting in reliable high efficiency thrusters. The concept, to be used in combination with advanced cryogenic storage technologies, is proposed for short time and high specific impulse orbit transfers. The simplified two-dimensional thermal design approach adopted is discussed, and its application to the engineering of the resistor is shown for both Inconel 718 and tungsten. The paper reports the performance testing of the proof-of-concept version of the thruster, manufactured in Inconel 718. Experiments on hydrogen show near ideal performance, demonstrating a peak thermal efficiency of 96%. The thruster proves the validity of the design methodology proposed, and the feasibility of the approach to develop monolithic additively manufactured hydrogen resistojets as main propulsion units.

The objective of this research is to study the impact of main engine parameters on the performance of a reusable liquid propellant rocket engine (LPRE). For this purpose, after a brief review over the available reusable liquid propellant rocket engines (RLPREs), the Space Shuttle Main Engine (SSME) is selected as a case study and a modular simulation software has been developed for analyzing the steady-state performance of LPRE in MATLAB Simulink. The evolution path of SSME is studied in detail and considering the reusability, two approaches have been selected. In the first approach, the performance of an expendable version of SSME is simulated and compared with the reusable version. It is found that the SSME working at 109% or higher rated thrust level (RTL) is the most possible option for the expendable version of this engine. The simulation results show that the components of the expendable SSME would work at 3–12% higher operating condition and the engine thrust would be 4.76% higher. In the second approach, in order to decrease the loads on the engine components and to enhance the reusability capability, the throat of the combustion chamber is enlarged. The simulation results show that by 11% increase in the combustion chamber throat area, the average load on the engine's different components would decrease 5–17% but for the same level of thrust, the specific impulse would be 2.43 s less. It is believed that the obtained results can help the designers to have a quantitative overview of the effect of reusability on the LPRE and launch vehicle performance during the conceptual design phase.

The aim of the present paper is the investigation of a monopropellant engine based on the exothermic decomposition of nitrous oxide, as a “green” substitute of the most common but highly toxic hydrazine-based systems. First, a general procedure for the design of such a system, in terms of sizing of the catalytic chamber and of the nozzle for prescribed mission requirements, is defined. This procedure is applied for the design of an 800-mN-class thruster prototype, with catalysts based on palladium as active phase on alumina pellets support. Such prototype is experimentally tested highlighting first the nitrous oxide decomposition behavior and issues related to an optimization of system thermal insulation for reducing energy losses. Finally, the measured engine performance are presented and discussed, including a comparison with the results obtained with the same system in a cold-flow operation mode, showing the gain in terms of increased specific impulse and reduced propellant consumption due to the exploitation of the thermal energy obtained from nitrous oxide decomposition.

The focus of this review is to collect and compare research studies on hybrid rocket fuels and components produced via additive manufacturing. The use of hybrid rocket propulsion for launching applications is limited due to its inherently low regression rate and actual specific impulse. To address these limitations, various additive manufacturing methods have been employed so far to fabricate fuel grains with complex port geometries or composite fuel grains possessing intricate shaped, printed scaffolds into which a second fuel material is incorporated. In addition to the fuel grains, additive manufacturing technology has also proven to be beneficial for fabricating a considerable number of hybrid rocket components, thereby reducing the part count, production time and cost. In this review, the rationale, types of the materials used, methods and comparison of performance data are presented. Future directions that use additive manufacturing to enhance hybrid rocket propulsion are also provided.

Electrospray thrusters with small size, high precision specific impulse, low power consumption are the most recommended propulsive equipments for micro-nano satellites. Molecular dynamics simulation is able to be used to predict the fundamental physics of electrospray and evaluate the propulsive performance. However, the accuracy of molecular dynamics simulation results critically depends on the accuracy of the interaction potential models used. Therefore, the present paper discusses the effect of different interaction potential models on the molecular dynamics simulations of electrospray thrusters with ionic liquid 1-ethyl-3-methyl-imidazolium tetrafluoroborate. The reduced charge all-atoms model, full charge all-atoms model, effective-force coarse-grained model and Merlet coarse-grained model reported in the literature are used for comparison, and comparison is also made to experimental data to elucidate which interaction potential model might be the most realistic. The process of Taylor cone formation, electrospray currents, energy characteristics and velocity distribution are analyzed under varying operating conditions. Our calculations indicate that the reduced charge all-atoms model is the most accurate model for molecular dynamics simulation of electrospray. In contrast, the coarse-grained models fail to reveal the energy characteristics. The full charge all-atoms model is not able to characterize the velocity characteristics. The full charge all-atoms model and coarse-grained models underestimate and overestimate the propulsive performance respectively. It is also found that the energy characteristics and velocity distribution of particles within the Taylor cone are almost unchanged with varying operating conditions.

A hollow cathode is an efficient electron source in the self-heating mode utilizing the discharge power. However, in sub-ampere currents, it needs keeper power to maintain the thermionic electron discharge, which could decrease the thrust efficiency. To address this problem, we propose using a microwave cathode, which is based on the flight model of a microwave ion thruster neutralizer cathode, as an alternative to a hollow cathode. First, we redesigned the magnetic field of a microwave cathode discharge chamber and tested it in the diode mode configuration. The electron emission current is doubled compared to the original performance. Next, we coupled the improved microwave cathode with a 200-W class Hall thruster and compared the characteristics and performance with a hollow cathode. We confirmed that the magnetic field polarity affects the ignition characteristics. We measured the thrust by an inverted thrust stand, the ion energy distribution functions by a retarding potential analyzer, and the beam profiles by an ion collector. The thrust and thrust efficiency are equivalent for both types of cathode. The specific impulse is 10% higher in the case of the microwave cathode. Since the potential difference between the microwave cathode and ground rapidly increased at currents above 600 mA, this could be taken to be the trade-off point against the hollow cathode., [Highlights] / • Electron current of a microwave cathode doubled by redesign of magnetic circuit. / • Magnetic polarity of microwave cathode and Hall thruster affected plasma maintenance. / • Thrust was equivalent in the both cases of microwave cathode and hollow cathode. / • Performance trade-off point between the two cathodes was about 600mA. / • Efficiency analysis by IEDF and beam profile showed different plasma characteristics.

Gas generator is an essential part of gas generator cycle Liquid Propellant Rocket Engine (LPRE). While designing the Gas Generator (GG), the selection of an unoptimized characteristic length result in an increase in mass flow rate of propellants required to drive turbine. As the exhaust of the gas generator is of little importance in the thrust of the engine, therefore, the specific impulse of the engine decreases with increase in the propellant flow rate to the gas generator. In literature, it is witnessed that researchers have conducted investigation of gas generator performance in terms of variation in mixing ratio or some models have been used to predict the characteristic length; however, it lacks the experimental investigation of gas generator performance in terms of characteristic length. In this paper, an experimental investigation of the gas generator of LPRE is conducted to determine the optimum characteristic length of the gas generator with liquid oxygen and kerosene. Hot tests are conducted with two different types of injectors, namely pressure swirl injector and impinging injectors in a modular water cooled gas generator. The gas generator is tested using six different characteristic lengths (i.e., from 1.75 m up to 9.67 m) for pressure swirl injector and with three different characteristic lengths (i.e., from 1.75 m up to 7.15 m) for impinging injector. The experimental results of hot tests show that the highest characteristic velocity of gas generator is achieved at a characteristic length of 4.27 m. It is demonstrated that a higher characteristic velocity of gas generator is achieved using impinging injector as compared to liquid-liquid pressure swirl coaxial injector at mixing ratio of 0.28 at optimum characteristic length. The effects of increase in mixing ratio with decrease in total mass flow rate on the chamber pressure are also discussed.

In this paper, the possibility of feedback servo control of a spacecraft thruster's specific impulse by a boundary feedback system is theoretically considered. The motivation to introduce feedback control is two-fold. The first is to stabilize any inherent plasma instabilities and the second is regulate the output specific impulse of the thruster. Two cases are considered: Electro-thermal thrusters and Electro-dynamic thrusters. The chamber inlet temperature in the case of the electro-thermal thruster or the boundary electric field potential responsible for generation of the velocity of the plasma ions in the case of an electrodynamic thruster, are controlled by feedback so as to regulate the thruster's specific impulse. By introducing typical disturbances in the plasma ionization voltage, it is shown using a two-dimensional fluid model and a suitable boundary feedback law, where the chamber inlet temperature or applied boundary potential is proportional to the error in the specific impulse and the desired specific impulse, that the specific impulse of the thruster may be regulated and held constant. The robustness of the control system is numerically tested, by a two dimensional simulation model using McCormacks's method. The Navier-Stokes equations, including the magneto-hydrodynamic variables were discretized and simulated, using the explicit MacCormack method for a typical nozzle domain. The numerical results for the open and closed-loop velocity fields were obtained and the specific impulse was computed from these fields. It was thus shown that not only the stability of the plasma is realized but also that the specific impulse is regulated as desired.

The use of krypton in the field of electric propulsion has become a key research topic owing to its high specific impulse and high reserves. However, because krypton has features such as a small ionization cross section and high ionization energy, its ionization performance is relatively low, thereby increasing the actual specific impulse loss of the krypton Hall thruster in comparison with its theoretical value. Therefore, to use of krypton as the propellant of the Hall thruster instead of xenon, it is necessary to study the optimization of its ionization process. In this study, a gas distributor with a rotating propellant supply mode is designed, considering factors such as the velocity and density of the velocity of neutral atoms, which affects the ionization process. The distribution of neutral atom parameters in the discharge channel is changed by the rotation; and its influences on the propellant ionization and other performance parameters of the Hall thruster are experimentally analyzed with respect to the magneto-ampere, volt-ampere, and flow rate-ampere characteristics. The results demonstrate that the rotating propellant supply reduces the axial velocity of neutral atoms, increases the density and circumferential motion distance, and thus improves the ionization efficiency of the krypton Hall thruster. Furthermore, the performance parameters, such as thrust and efficiency, are also improved, and the plume divergence angle is reduced. When the discharge voltage is 450 V and the anode flow rate is 60 sccm, the rotating propellant supply can increase the anode efficiency of the axial propellant supply mode from 46.2% to 53.2%. However, the plume divergence half-angle reduces from 39.2° to 35.5°. In addition, an increase in the anode flow rate and discharge voltage continuously increases the performance promotion in the rotating propellant supply mode.

This paper presents the design of a low power (8 W) water-fed vaporizing liquid microthruster that operates in the Leidenfrost boiling regime. The direct thrust measurements show the proposed design is able to produce one of the best specific impulse amongst all similar microthrusters reported in the literature. One of the important features that enables the superior performance is the high heat flux miniature molybdenum heating elements fabricated inside the vaporization chamber in a single silicon device package of dimensions 35 mm ✕ 35 mm ✕ 1.05 mm. The molybdenum heating elements are capable of producing a heat flux output of 370 kW/m2 resulting in a sustained operational pressure of 2 bar and temperature of 446 K. The specific design details, such as choice of configurations and sizing, are given in this paper. In particular, the choice of material and the fabrication process carried out at the in-house nanosystem fabrication facility are described. The experimental setup and uncertainty analysis of the mN-range hanging pendulum direct thrust measurement stand are presented. Then, the evaporation of liquid water propellant is investigated in detail. The occurrence of Leidenfrost effect inside the vaporization chamber is captured with a high-speed camera at a frame rate of 5000 Hz. It can be seen that the rapid vaporization of propellant in the Leidenfrost boiling regime results in a high specific impulse amongst other microelectromechanical-system-scale vaporizing liquid microthrusters. The direct thrust measurements show the current design is capable of producing a specific impulse ranging between 25.9–51.1 s at a continuous thrust level of 1.00–2.00 mN; considering contemporary energy storage technologies, the estimated thrust per device dry mass and specific impulse per device dry mass levels are 4.0–8.0 mN/kg and 103.6–204.0 s/kg respectively.

Cyclotrimethylene trinitramine (RDX) is widely used in aluminized solid propellants to increase specific impulse. Experimental investigations on the effect of RDX content on the propellant combustion and agglomeration were thoroughly conducted based on Thermogravimetry-differential scanning calorimetry, laser ignition, combustion diagnosis and a new home-made condensed combustion products (CCPs) collection device. The results show RDX inclusion inhibits the decomposition of AP. The ignition delay and self-sustaining combustion time of aluminized propellant increase, while the combustion intensity decreases with increasing RDX content. RDX addition is found to decrease the burning rate within 6–10 MPa. Compared with the baseline propellant, RDX leads to obvious aggravation of aluminum agglomeration on the propellant burning surface. The 6 wt% and 12 wt% RDX-included propellants increase the average CCPs size from 46.3 μm to 86.7 μm and 96.6 μm, respectively. The combustion efficiency of aluminum in propellant is reduced by 15% when the RDX content is increased from 0 to 12%. Overall, RDX plays a comprehensive effect on the ignition, combustion and agglomeration characteristics of aluminized propellant. Results of this study can help guide the application and development of RDX in aluminized propellants for solid rocket motors.

Using high power space transportation systems equipped with electric propulsion requires development and qualification of electric thrusters with power equal to tens of kW. The optimal specific impulse of such systems designed for deep space missions should be very high (6000…8000 s). Elements for electric propulsion systems are developing at Keldysh Research Center (KeRC). This paper presents qualification status of IT-500 35-kW ion thruster and FCU-500 flow control unit. 2000-hours lifetime test of the IT-500 and FCU-500 was performed as a part of qualification. The ion thruster accomplished 2018 hours of operation for the most part at input power of 17,8 kW, utilized 40 kg of xenon. Magnetic field and ion optics improvement and graphite grids development status are also described in this paper.

This paper presents the first proof of concept validation of the STAR thruster prototype. The device contains an innovative multifunctional monolithic heat exchanger, enabled by metal additive manufacturing processes. A 316L stainless steel printed thruster is characterized through a combination of dry heating and wet firing tests. This includes verification testing with argon in both cold and hot firing mode, at a range of electrical power inputs. Thrust measurements range from 9.7 mN ± 0.16 mN–29.8 mN ± 0.16 mN, with a maximum measured specific impulse of 80.11 ± 1.49 s. Thrust performance is measured using a high-precision balance, and liquid-metal power transfer terminals to eliminate thermal drift. Highly coupled multiphysics computational models provide validation of the electro-thermal and thermo-fluidic characteristics of the prototype, including a prediction of the maximum propellant stagnation temperature and structural temperature, which were 649 °C and 854 °C.

An 85 W class Hall thruster with inner and outer electromagnetic coils was developed, and its thrust performance and plume characteristics were experimentally evaluated. The 85-W Hall thruster required a maximum magnetic flux density along the channel centerline above 24.3 mT to achieve stable and efficient operation above a discharge voltage of 225 V. A specific impulse of 1050 s, thrust-to-power ratio of 59.5 mN/kW, and anode efficiency of 0.306 were achieved at a discharge voltage of 225 V and a discharge power of 88.9 W under a background pressure of 7.9 × 10−4 Pa. Moreover, the thruster achieved throttling at a discharge power of 47.6–118.5 W with stable operation. The low anode efficiency of the 85-W Hall thruster was due to low propellant utilization and a large beam divergence. The short effective length for the propellant ionization owing to the narrow channel is deduced to be the cause for the low anode efficiency in low-power Hall thrusters. The mass utilization efficiency of the 85-W Hall thruster improved as the anode mass flow density increased. Therefore, an increase in the anode mass flow density is required to improve the performance of the 85-W Hall thruster.

This study proposes an embedded rocket ramjet, a novel combined ramjet engine, that can operate at high altitude and low pressure. The embedded rocket ramjet combustor is investigated experimentally and numerically to predict its characteristics. Experiments are performed in a direct-connect ground test facility to measures the wall pressure and total pressure at the combustor exit. Three-dimensional Navier–Stokes equations and species transport equations are solved numerically. A compressible reactive shear stress transport k-ω turbulence model with reduced chemical kinetics of a surrogate model for kerosene is used to study the flow characteristics and overall performance of the combustor. Test results show that the embedded rocket can significantly increase the combustor pressure. Numerical analysis shows that the embedded rocket increases the mixing efficiency, combustion efficiency, and thrust but decreases the specific impulse.

A rocket-based combined-cycle (RBCC) engine synthetically integrates a traditional rocket engine with an airbreathing engine, effectively combining both advantages of high thrust-to-weight ratio and high specific impulse. It has become one of the most efficient propulsion systems for future reusable space transportation vehicles. In engineering practice, the engine performance during ejector mode plays a significant role in determining the efficiency of the whole RBCC propulsion system. This study aims to provide a summary report on the recent research progress on the RBCC engine in ejector mode. Firstly, the basic operating principle of the RBCC ejector mode is introduced and its ideal thermodynamic cycle process is theoretically analyzed. Secondly, the influencing factors on RBCC engine performance in ejector mode are specifically investigated, including rocket ejector parameters, geometric configurations of RBCC engine, secondary fuel injection schemes, flight Mach number and altitude. Thirdly, the research progress on key techniques of the RBCC ejector mode in recent two decades is reviewed from four aspects, namely the combustion organization, the mixing enhancement, the resistance to backpressure and the ejector/ramjet mode transition. Finally, some suggestions for the future work on the RBCC ejector mode are proposed.

Evaluating the climbing performance of rocket-based combined cycle (RBCC) powered aircraft has a guiding role for single-stage to orbit. This study establishes a new modularized thrust model of the RBCC engine for low speeds (Ma∞ = 0.8–3) based on independent ramjet stream and pressure matching. The model can quickly estimate the quasi one-dimensional flow characteristics of secondary flow and primary flow at given runner geometry and operating conditions. Comparison with experimental results shows that the thrust model is correct. In this model, the influence of altitude, Mach number, and rocket flow on thrust and specific impulse is considered. The model was used to optimize the climbing trajectory of RBCC powered aircraft with a typical lifting-body. The comparison of RBCC propulsion and pure rocket propulsion shows that the former offers propellant saving advantages. During the climb, the RBCC powered aircraft needs to strictly control the air flow rate and ensure the pressure matching in the flow path to improve propulsion efficiency.

The high specific impulse of the ramjet engine, coupled with the potential of a gel to carry boron particles, makes the boron-based gel fuel ramjet a most appropriate solution for a mid-high range sustainer. The current research work aims at verifying the empirical feasibility of this kind of a concept, besides studying the combustor performance. A test facility and a lab-scale motor have been designed and developed for the investigation of the impact of the chamber length and boron particle content on the combustion efficiency of the motor. As the results suggest, the combustion efficiency declines with the increasing boron particle content. Moreover, a higher combustion efficiency was attained with the longer chamber. As the mass fraction of boron particles reaches 40%, the combustion efficiency, which is based on the temperature rise, still has the potential to reach 80%.

REGULUS is a propulsion platform for CubeSats. It integrates the Magnetically Enhanced Thruster (MET) and its subsystems (i.e., fluidic line, electronics, and thermo-structural components) in a 2U envelope of weight lower than 3 kg. MET is a RF cathode-less thruster capable of providing thrust in the range 300 μN to 900 μN, with maximum specific impulse of 900 s and maximum operation power of 60 W. The performance has been measured with a dedicated thrust balance, when the thruster is operated with 0.1 mg/s of Xenon gas. REGULUS: (i) is compatible with thermal, mechanical and electrical interfaces of the CubeSat platforms available in the marked, (ii) it relies on COTS components to abate the recurrent costs, and (iii) it is based on a passive thermal control system. The main framework of application of the REGULUS platform is the propulsion of medium-to-large CubeSats (from 6U up to 27U). Such capability is here demonstrated by comparing MET with other thrusters for CubeSats offered in the market. The feasibility of two possible mission scenarios, namely drag compensation and constellation deployment, has been verified through preliminary orbit calculations. The REGULUS platform can compensate the effects of atmospheric drag on a 6U CubeSat in a 400 km altitude orbit for years. Besides, by exploiting the natural drift of the ascending node caused by the second zonal harmonic of the terrestrial gravity field, a CubeSat constellation can be deployed through several planes in some months employing small fractions of the onboard propellant.

Fundamental characteristics of a microwave-discharge miniature ion thruster using water as a propellant, instead of the conventional xenon, are presented in this paper intended for future applications for CubeSats. Water has advantages in terms of safety, handling ability, and availability compared to any other propellant, and these features are especially important for CubeSats. The ion thruster consists of an ion source and a neutralizer, and both use an electron cyclotron resonance heating to generate the plasma. Water was fed to the ion thruster or the neutralizer as a vapor and the direct thrust measurement was conducted. As a result, the thrust correction factor of the ion source was 0.92 ± 0.074 with a beam acceleration voltage of over 0.80 kV, which was similar to the typical values for xenon ion thruster. The ion beam was extracted at the lower mass flow rate, the highest thrust was 164 ± 13.4 μN, and the highest specific impulse was 665 ± 59 s while neglecting the neutralizer. The electron emission current from the neutralizer was 9.5 ± 0.08 mA, which is sufficient to neutralize the ion beam. These experiments demonstrated the basic features of the miniature ion thruster using water and a future possibility of water propellant utilization for CubeSats.

Accepted: 2018-12-17, 資料番号: SA1180298000

In recent years, satellite design has extended towards miniaturisation to reduce associated cost with launching and conducting space missions. Small satellites provide low-cost platforms for space missions. However, this lower cost comes at the expense of the removal of key sub-systems, such as the propulsion system, due to the small available onboard volume and mass restrictions. For this reason, small, lightweight, high-performing and affordable propulsion systems are necessary. However, there is limited research available on the comparison of propulsion technologies with regards to cost. Motivated by the above challenges the objective of this paper is to provide a comparison of propulsion technologies that are compatible with small satellites with respect to cost and application. The different propulsion systems are investigated for three mission scenarios, a small on-orbit manoeuvre, a station-keeping, and a lunar orbit transfer mission. Each system is evaluated in terms of a total figure of merit which incorporates nine variables such as propellant mass, safety, and hardware price, that affect the total cost of a propulsion system. This figure of merit is used to quantitatively compare the propulsion systems to identify cost-effective solutions as a function of various mission scenarios. Solar thermal propulsion has been proposed for small satellite applications, but information regarding the concepts are not available in a single report. Therefore, a secondary objective of this paper is to provide the reader with a review of the current status of solar thermal propulsion. An important finding of this research is the classification of propulsion systems in terms of thrust, specific impulse, cost, and application.

This work addresses the growing need for an intuitive, systematic approach to low-thrust spacecraft formation flying by extending shape-based continuous thrust trajectory design methods to the relative motion of two spacecraft. There is widespread interest in distributed space systems for their low costs, broad capabilities, and high redundancy. This trend introduces a new challenge for trajectory design when combined with the increasing prevalence of low-thrust, high specific impulse electric propulsion systems. That challenge is met herein with a geometrically intuitive, semi-analytical solution to the low-thrust problem. Beginning with the equations of relative motion of two spacecraft, an unperturbed chief and a continuously-thrusting deputy, a thrust profile is constructed which transforms the equations into a form that is solved analytically. The resulting relative trajectories are the family of sinusoidal spirals, which provide diversity for design and optimization based upon a single thrust parameter. Closed-form expressions are derived for the trajectory shape and time-of-flight for two prescribed relative velocity behaviors, and used to develop a novel patched-spirals trajectory design and optimization method. The example problem of a servicer spacecraft establishing and reconfiguring a formation around a target in geostationary earth orbit is used to demonstrate the application of the patched spirals technique as well as the advantages of the relative spiral trajectories over impulsive maneuvers. The sensitivity of the trajectory solutions to deviations from the underlying assumptions, uncertainties in the state, and errors in thrust are studied through high-fidelity simulation.

The use of a variable geometry combustor is one of the most effective methods to improve the performance of a rocket-based combined-cycle (RBCC) engine over a wide range of operating conditions. This paper aims to study the capabilities of a variable geometry combustor operating over a wide range of conditions and determine the performance of the combustor under various inflow conditions. Based on the inflow conditions, the configuration parameters of the combustor were adjusted. Ground direct-connect experiments were conducted under the inflow conditions of Ma 2, Ma 3, Ma 4, and Ma 6, and numerical simulations were performed under the conditions of Ma 3 and Ma 6. The direct-connect experiments showed that the primary rocket operating at a low flow rate can reliably ignite the secondary fuel and maintain a stable and efficient combustion in a variable geometry combustor. Under the inflow conditions of both Ma 4 and Ma 6, smooth transitions of the variable geometry combustor from rocket-ramjet mode to ramjet mode were achieved, and the specific impulses were greatly improved and reached 28.2% and 37.1% of the amplitude, respectively. The numerical simulation showed that the variable geometry combustor can effectively control the combustion heat release region and greatly improve the performance of the combustor. The specific impulse in the combustor increased by 18.6% and 26.2% compared with that in the fixed geometry combustor under Ma 3 and Ma 6 inflow conditions, respectively. It is therefore strongly believed that the variable geometry combustor has significant performance advantages under a wide range of operating conditions.

Estimation of a possibility to use an electromagnetic plasma railgun as a safe and reliable plasma thruster based on calculations of the thrust force and specific impulse of a plasma jet flowing from an accelerator channel is presented. Theoretical and experimental investigations of acceleration of a plasma armature in railguns filled with inert gases (argon, xenon) carried out at Ioffe Institute are described.

Hydrogen fuel, as for now, is the most energy efficient and clean fuel for aviation engines. At present time RAM engines based on chemical burning of fuels and accelerating exhaust gases practically reached the top of their efficiency in terms of specific impulse. New schemes are under development. As it is known, complex geometry of combustion chambers could bring to compression waves focusing and changing the mode of flame propagation from slow combustion to detonation. Keeping this process under control is of major importance for developing new generation of engines. The paper presents results of numerical and experimental investigation of mixture ignition and detonation onset in shock wave reflected from inside a wedge. Comparison of numerical and experimental results made it possible to validate the developed 3-D transient mathematical model of chemically reacting gas mixture flows incorporating hydrogen – air mixtures. Kinetic schemes were improved based on comparison of numerical and experimental results. A tool was developed making it possible to set a criterion for obtaining either safe combustion regime, or transition to detonation regime in shock wave focusing.

Solid-fuel ramjet converts thermal energy of combustion products to a forward thrust without using any moving parts. Normally, it uses air intake system to compress the incoming air without swirler. A new design of swirler has been proposed and used in the current work. In this paper, a series of firing tests have been carried out to investigate the impact of using swirl flow on regression rate, combustion characteristics, and performance of solid-fuel ramjet engines. The influences of swirl intensity, solid fuel port diameter, and combustor length were studied and varied independently. A new technique for determining the time and space averaged regression rate of high-density polyethylene solid fuel surface after experiments has been proposed based on the laser scan technique. A code has been developed to reconstruct the data from the scanner and then used to obtain the three-dimensional distribution of the regression rate. It is shown that increasing swirl number increases regression rate, thrust, and characteristic velocity, and, decreases air-fuel ratio, corner recirculation zone length, and specific impulse. Using swirl flow enhances the flame stability meanwhile negatively affected on ignition process and specific impulse. Although a significant reduction of combustion chamber length can be achieved when swirl flow is used. Power fitting correlation for average regression rate was developed taking into account the influence of swirl number. Furthermore, varying port diameter and combustor length were found to have influences on regression rate, combustion characteristics and performance of solid-fuel ramjet.

This study proposes a turbocharged solid propellant ramjet (TSPR) propulsion system that integrates a turbocharged system consisting of a solid propellant (SP) air turbo rocket (ATR) and the fuel-rich gas generator of a solid propellant ramjet (SPR). First, a suitable propellant scheme was determined for the TSPR. A solid hydrocarbon propellant is used to generate gas for driving the turbine, and a boron-based fuel-rich propellant is used to provide fuel-rich gas to the afterburner. An appropriate TSPR structure was also determined. The TSPR's thermodynamic cycle was analysed to prove its theoretical feasibility. The results showed that the TSPR's specific cycle power was larger than those of SP-ATR and SPR and thermal efficiency was slightly less than that of SP-ATR. Overall, TSPR showed optimal performance in a wide flight envelope. The specific impulses and specific thrusts of TSPR, SP-ATR, and SPR in the flight envelope were calculated and compared. TSPR's flight envelope roughly overlapped that of SP-ATR, its specific impulse was larger than that of SP-ATR, and its specific thrust was larger than those of SP-ATR and SPR. Attempts to improve the TSPR off-design performance prompted our proposal of a control plan for off-design codes in which both the turbocharger corrected speed and combustor excess gas coefficient are kept constant. An off-design performance model was established by analysing the TSPR working process. We concluded that TSPR with a constant corrected speed had wider flight envelope, higher thrust, and higher specific impulse than TSPR with a constant physical speed determined by calculating the performance of off-design TSPR codes under different control plans. The results of this study can provide a reference for further studies on TSPRs.

Due to the current global energy shortage and increasingly serious environmental issues, green propellants are attracting more attention. In particular, the ammonium dinitramide (ADN)-based monopropellant thruster is gaining world-wide attention as a green, non-polluting and high specific impulse propellant. Gel propellants combine the advantages of liquid and solid propellants, and are becoming popular in the field of spaceflight. In this paper, a swirling atomization experimental study was carried out using an ADN aqueous gel propellant under different injection pressures. A high-speed camera and a Malvern laser particle size analyzer were used to study the spray process. The flow coefficient, cone angle of swirl atomizing spray, breakup length of spray membrane, and droplet size distribution were analyzed. Furthermore, the effects of different injection pressures on the swirling atomization characteristics were studied.

CubeSats have been extensively used in the past decade as scientific tools, technology demonstrators and for education. Recently, PocketQubes have emerged as an interesting and even smaller alternative to CubeSats. However, both satellite types often lack some key capabilities, such as micropropulsion, in order to further extend the range of applications of these small satellites. This paper reviews the current development status of micropropulsion systems fabricated with MEMS (micro electro-mechanical systems) and silicon technology intended to be used in CubeSat or PocketQube missions and compares different technologies with respect to performance parameters such as thrust, specific impulse, and power as well as in terms of operational complexity. More than 30 different devices are analyzed and divided into 7 main categories according to the working principle. A specific outcome of the research is the identification of the current status of MEMS technologies for micropropulsion including key opportunities and challenges.

The sequential use of an electric propulsion system is considered in combination with a high-thrust propulsion system for application to the propellant-optimal Geostationary Orbit insertion problem, whilst considering both temporal and radiation flux constraints. Such usage is found to offer a combined propellant mass saving when compared with an equivalent high-thrust only transfer. This propellant mass saving is seen to increase as the allowable transfer duration is increased, and as the thrust from the low-thrust system is increased, assuming constant specific impulse. It was found that the required plane change maneuver is most propellant-efficiently performed by the high-thrust system. The propellant optimal trajectory incurs a significantly increased electron flux when compared to an equivalent high-thrust only transfer. However, the electron flux can be reduced to a similar order of magnitude by increasing the high-thrust propellant consumption, whilst still delivering an improved mass fraction.

The objective of this effort is to study the combustion performance of a hybrid rocket motor with the help of 3D steady-state numerical simulation, which applies 90% hydrogen peroxide as the oxidizer and hydroxyl-terminated polybutadiene as the fuel. A method of secondary oxidizer injection in post-chamber is introduced to investigate the flow field characteristics and combustion efficiency. The secondary injection medium is the mixed gas coming from liquid hydrogen peroxide catalytic decomposition. The secondary injectors are uniformly set along the circumferential direction of the post-chamber. The simulation results obtained by above model are verified by experimental data. Three influencing parameters are considered: secondary injection diameter, secondary injection angle and secondary injection numbers. Simulation results reveals that this design could improve the combustion efficiency with respect to the same motor without secondary injection. Besides, the secondary injection almost has no effect on the regression rate and fuel sueface temperature distribution. It is also presented that the oxidizer is injected by 8 secondary injectors with a diameter of 7–8 mm in the direction of 120°in post-chamber is identified as the optimized secondary injection pattern, through which combustion efficiency, specific impulse efficiency as well as utilization of propellants are all improved obviously.