Space Propulsion Systems Research Articles

SiC nanograins stabilized Si–C–B–N fibers with ultrahigh‐temperature resistance

AbstractContinuous ceramic fibers with ultrahigh‐temperature stability are in high demand for applications in advanced space propulsion and thermal protection systems. In this study, SiC nanograins stabilized Si–C–B–N ceramic fibers were prepared using chemically modified polyborosilazane via a polymer‐derived method. The fabricated Si–C–B–N fibers exhibited a rather high tensile strength of approximately 1.8 GPa and a high strength retention of approximately 90% after annealing at 2100°C for 0.5 h under a nitrogen atmosphere. The ultrahigh‐temperature stability can be contributed to the presence of thermodynamically stable SiC nanograins and the encapsulation of SiC nanograins by the BN(C) phase and amorphous Si–C–B–N matrix. Our work offers a convenient strategy for preparing Si‐based ceramic fibers with ultrahigh‐temperature stability at beyond 2000°C.

The approach to numerical simulation of the spatial movement of fluid with forming free gas inclusions in propellant tank at space flight conditions

The space propulsion systems ensure se veral start-ups and shutdowns of main liquid-propellant rocket engines under microgravity conditions for the spacecraft program movements and reorientation control. During the passive flight of the space stage (after its main engine shutdown), the liquid propellant in the tanks continues to move by inertia in microgravity away from the propellant management device as much as possible. In this case, the pressurization gas is displaced to the propellant management device, which creates the potential danger of gas entering the engine inlet in quantities unacceptable for the reliable engine restart. In this regard, determining the parameters of fluid movement in propellant tanks in microgravity conditions is an urgent problem that needs to be solved in the design period of liquid propulsion systems. We have developed an approach to the theoretical computation of the parameters of the motion of the ‘gas — fluid’ system in the propellant tanks of modern space stages in microgravity conditions. The approach is based on the use of the finite element method, the Volume of Fluid method and modern computer tools for finite-element analysis (Computer Aided Engineering — CAE systems). For the passive leg of the launch vehicle space flight, we performed mathematical modeling of the spatial movement of liquid propellant and forming free gas inclusions and determined the parameters of movement and shape of the free surface of the liquid in the tank as well as the location of gas inclusions. The numerical simulation of the fluid movement in an experimental sample of a spherical shape tank was performed with regard to the movement conditions in the SE Yuzhnoye Design Bureau ‘Drop tower’ for studying space object s in microgravity. The motion parameters of the ‘gas — fluid’ interface obtained as a result of mathematical modeling are in satisfactory agreement with the experimental data obtained. The use of the developed approach will significantly reduce the amount of experimental testing of the designed space stages.

Interstellar Propulsion Using Laser-Driven Inertial Confinement Fusion Physics

To transport a spacecraft to distances far beyond the solar heliosphere and around the planets of other stars will require advanced space propulsion systems that go beyond the existing technological state of the art. The release of fusion energy from the interaction of two low mass atomic nuclei that are able to overcome the Coulomb barrier offers the potential for ∼1011J/g specific energy release and implies that robotic missions to the nearby stars to distances of ∼5–10 ly may be possible in trip durations of the order of ∼50–100 years, travelling at cruise speeds of the order of ∼0.05–0.15 c. Such missions would be characterised with ∼kN-MN thrust levels, ∼GW-TW jet powers, ∼kW/kg-MW/kg specific powers. One of the innovative methods by which fusion reactions can be ignited is via the impingement of laser beams onto an inertial confinement fusion capsule, imploding it to a thermonuclear state. This paper gives an overview of the physics of inertial confinement fusion and the interaction of a laser beam with a capsule to include the simulation of a 1D particle-in-cell code calculation to illustrate the effects. In the application to deep space missions, various spacecraft concepts from the literature are discussed, and the range of values assumed for the pulse frequency, burn fraction and areal density appropriate for the mission are presented. It is concluded that advanced space propulsion via inertial confinement fusion is a plausible part of our future, provided that experimental validation of ignition is on the horizon and numerical models for feasibility concepts are developed to high fidelity and on a consistent basis.

Helicon Injected Inertial Plasma Electrostatic Rocket

The Helicon Injected Inertial Plasma Electrostatic Rocket (HIIPER) is a space propulsion system developed at the University of Illinois Urbana-Champaign. The HIIPER couples a helicon tube with an inertial electrostatic confinement (IEC) fusion system. Its operating principle involves a helicon ionization stage followed by an electrostatic grid (IEC cathode grid) extraction stage. The helicon setup used in the HIIPER is modified to include a helicon bias grid at the upstream end of the tube. This grid is applied with a positive direct-current voltage to increase the plasma potential and the most probable ion energy of the plasma injected into the IEC fusion chamber. The IEC cathode grid in the HIIPER uses an innovative asymmetric design, graphically depicted through a computational model, that ejects a stream of electrons that accelerate the exhaust ions and simultaneously neutralize the exhaust jet. The model is also used to plot ion trajectories inside the HIIPER to identify any wall collision losses. A separate numerical study was undertaken to show augmentation of plasma kinetic energy on adding a magnetic nozzle as the final propulsion stage of the HIIPER. Experimental results were used to establish a relation between the input parameters and the ion density of the resulting plasma. Langmuir probe measurements were performed at two locations to validate corresponding computational results, indicating ion losses due to ion-wall collisions inside the helicon-IEC coupling. The results in this study add to the proof of concept of the HIIPER and allow for designing an upgrade of the propulsion system. Increasing thrust while maintaining plasma densities between 1017 and 1018 throughout the system is the current aim of HIIPER research. This study summarizes the various performance parameters of the propulsion system, along with a discussion of ongoing research and future scope.

Numerical Study of Effect of Pressurant Gas Species on Thermal Behavior in Cryogenic Tank

This study numerically reveals effect of pressurant gas species on thermal behavior in a cryogenic propellant tank for space propulsion systems. The simulation was conducted for a ground experiment using liquid nitrogen in which gaseous helium and the same species as the liquid, gaseous nitrogen, were used as prepressurants. The tank was sealed after prepressurization to observe self-pressurization (pressure rise with evaporation) and thermal stratification (high-temperature liquid near the surface). A higher evaporation rate and a thicker thermal layer were observed for helium prepressurization in the experiment. To clarify the underlying physics, a simulation considering phase change and conjugate heat transfer was developed. The simulated pressure and temperature closely matched the experimental results. Conjugate heat transfer played an important role; that is, the heat flow through the ullage part of the tank skin into the cryogenic liquid was a dominant heat transfer process, causing evaporation to occur mainly at the contact point of the wall and the gas–liquid interface. Afterward, nitrogen vapor rose along the tank wall due to buoyancy under nitrogen prepressurization. However, buoyancy in the ullage is lower under helium prepressurization since light helium stayed in the upper part of the ullage, and a radial vapor flow is produced from the contact point, leading to a higher heat flux from the tank wall to the liquid nitrogen. This mechanism shows that evaporation rate and heat flow into the liquid are higher for helium prepressurization than for nitrogen prepressurization.

A dual-mode hydrogen peroxide and kerosene space propulsion system pressurized by electric pumps

In this study, the performance characteristics and previous application cases of the dual-mode propulsion system (DMPS) were discussed. Besides, a dual-mode hydrogen peroxide and kerosene space propulsion system pressurized by electric pumps was proposed via the analysis of the existing problems with respect to the dual-mode propulsion system. In addition, the working process of this system was analyzed by system modeling and simulation. Further, the feasibility of DMPS was verified by exploring the working characteristics of the bipropellant main thruster and the monopropellant attitude control thruster. As per the analysis results, the DMPS proposed in this study contributed to eliminating the problems in the application of the dual-mode propulsion system at the present stage, and it also provided effective guidance in improving the performance of the propulsion system, adjusting the thrust flow rate, and making the propellant non-toxic.

TEST BENCH EQUIPMENT OF KELDYSH RESEARCH CENTRE FOR TESTING HIGH-PERFORMANCE POWER AND PROPULSION SYSTEMS

Expanding space exploration in near-Earth space and deep space calls for an increase in power supply of space vehicles to a level of hundreds of kilowatts and higher. Pre-requisite for successful implementation of works on development of power propulsion systems of such power level, including those with nuclear energy sources, is availability of a full-scale experimental test bench equipment. A unique experimental test bench has been built and is successfully operating at JSC «Keldysh Research Center», which allows performing both research and developmental tests of key components of high-performance power and propulsion systems under conditions close to nominal operating conditions. The article describes the main components of JSC «Keldysh Research Center» test bench equipment, presents some results of the work fulfilled on it, outlines prospects for its furtherance and use in development of space high-performance power and propulsion systems. Key words: testing, ground developmental testing, power conversion system, test bench equipment, electric propulsion, power and propulsion system.

Mutually guided light and particle beam propagation

The polarizability of atoms and molecules gives rise to optical forces that trap particles and a refractive index that guides light beams, potentially leading to a self-guided laser and particle beam propagation. In this paper, the mutual interactions between an expanding particle beam and a diffracting light beam are investigated using an axisymmetric particle-light coupled simulation. The nonlinear coupling between particles and photons is dependent on the particle beam radius, particle density, particle velocity and temperature, polarizability, light beam waist, light frequency (with respect to the resonance frequency), and light intensity. The computational results show that the maximum propagation distance is achieved when the waveguiding effect is optimized to single-mode operation. The application of the coupled beam propagation as a space propulsion system is discussed.

Development of metallic transition joints for space propulsion systems applications

Development of metallic transition joints for space propulsion systems applications

Effect of Xe plasma processing on characteristics of carbon nanotube-based field emission cathodes

We developed an experimental setup and procedure to investigate improving the electron emission stability in the long-term operation of carbon nanotube (CNT)-based field emission cathodes (FECs) by using Xe plasma processing aging method. To find the optimum plasma processing conditions, we varied the FEC's exposure time, plasma environment, and bias voltage. Plasma processing increased the gate voltage required for a certain number of electron emissions, whereas we confirmed that the emission stability improved and the aging effect of plasma processing. A bias voltage of approximately −15 V and a shorter exposure time were most suitable for plasma processing. Because the number of ion and electron collisions and the collision energies cause a moderate change in the CNT structure with minimal damage to the CNTs, we expect that the proposed plasma processing will improve the lifetime of CNT-based FECs, which is essential if they are to be used as electron emission sources for space propulsion systems.

Comparative study of plume characteristics of polymeric propellants filled with different oxides for laser-electric hybrid thruster

Laser-electric hybrid thrusters have attracted increasing attention as a new concept for space propulsion systems because of their low weight, high propulsion performance, and controllable thrust. When the thruster works, the laser-ablation plasma induced by the laser irradiation of a solid target is additionally accelerated by electrical means in the discharge chamber, generating thrust. Therefore, a stable and continuous ablation plume is required to ensure the subsequent electromagnetic acceleration process. This study compared and analyzed the plume characteristics of polymeric propellants doped with different oxides under the same operating conditions. The plume images were examined using an ultrahigh-speed camera, and the plume length, expanding velocity, plume area, and normalized intensity distribution were collected from the plume images based on the image processing method. The results illustrated that the CuO dopant contributed to the rapid, stable, and massive formation of the ablation plume, and VO2 could make the plume have a certain distance from the ablation surface, and then reduce the particle reflux deposition effect to a certain extent. It was also found that the samples doped with CuO, NiO, and Fe2O3 generated a micro cluster with a duration of about 0.5 ms at the beginning of ablation, while the sample doped with VO2 did not generate such a cluster. There were two regions of high-intensity light in the plumes of the samples doped with CuO and VO2. Furthermore, the plume areas of the samples doped with NiO, VO2, and Fe2O3 experienced a quiescence of approximately 2 ms before rapid growth, while that of CuO continued to increase steadily.

Combustion of Gelled HAN/Methanol/Water Propellants

AbstractMonopropellants based on hydroxylammonium nitrate (HAN) are a low‐toxic, high‐performance alternative to hydrazine in space propulsion systems. Gelling such propellants may offer additional advantages such as the long‐term stability of compositions with added energetic and catalytic particles in suspension. However, the effect of gelling on the combustion characteristics and mechanisms of HAN‐based propellants is not well understood. In the present work, an aqueous HAN/methanol solution (70.1 wt% HAN and 14.9 wt% methanol) was gelled with 1 wt% polyacrylamide in an acoustic mixer. Combustion of the obtained gel was studied in a strand burner at nitrogen pressures up to 30 MPa. Thermogravimetric analysis, differential scanning calorimetry, and mass spectrometry were used to investigate the reaction mechanisms in the HAN/methanol/water solutions with and without polyacrylamide. The strand burner tests have shown that gelling suppressed the multiple pressure regimes observed in previous studies of the HAN‐based propellants, with a consistent pressure exponent of 1.15±0.05 over the tested pressure range 4–30 MPa. Thermoanalytical testing of the gelled mixture indicates there is no chemical interaction between the gellant and the HAN decomposition, and the effect of gelling on the combustion of HAN/methanol/water propellants is hydrodynamic in nature.

Thermal Analysis and Testing of Different Designs of LaB6 Hollow Cathodes to Be Used in Electric Propulsion Applications

LaB6 emitters are commonly used in hollow cathodes that are utilized in electric space propulsion systems. In order to obtain necessary surface current emission densities of 1–10 A/cm2 for cathode operations, LaB6 emitters require temperatures above 1500 °C. Hence, the design for LaB6 cathodes presents thermal and mechanical challenges. In this paper, several design iterations for LaB6 hollow cathodes are presented and thermal analyses are conducted for each design. Temperature and heat flux distributions that are obtained from thermal analyses are investigated. The designs are evaluated according to the required heat input to the emitter, and the radiative and conductive heat loss mechanisms. In addition to the thermal analyses, experimental tests are conducted for different cathode designs and, based on these tests, various modes of failure are determined. Revising the cathode design and the material selection iteratively to eliminate the encountered failure mechanisms, a novel cathode design is achieved. Experimental tests of this novel cathode are conducted and current-voltage characteristics are presented for different mass flow rates and for discharge currents between 0.5 and 12 A. Tests and analysis results show that the selection of materials and design are crucial for a sturdy and long lifetime cathode.

Experimental Investigation of configuration types and array of external heat sources on downward spreading fires

Scholars and scientists have been attempting to discover ways to control and lessen the reverberation of concurrent fires such as forest fires, building fires, and other various space fires but no convincing solutions have been concluded from their studies so far. The fundamental genesis of these types of fires concerns the unstable nature of the flames and considerable unpredictability associated with them. This led us to make an effort to study the etiquette of such flames – an experimental setup with rather ideal conditions was devised and an in-depth study was carried out. The present study predominately covers the study of fire propagation phenomena and the zeal of fire control in our daily life. This study includes the review of the varying regression rates and fire spread rates of flames as noticed in matchsticks when spaced in a linear orientation. The deportment instability of the flames will give us an insight into the heterogeneous fire propagation phenomenon and its control. Intuition into the heterogeneous fire propagation is expected for essential fire safety and, on its basis, an algorithm for the same is to be formulated. With this understanding along with the existing information, it might give us some possible solution to the reduction of such kinds of fires in buildings, forests, space propulsion systems and, large-scale fires in industries.

Space Nuclear Power and Propulsion at USNC-Tech

USNC-Technologies (USNC-Tech) is an advanced nuclear company focused on developing space nuclear power and propulsion systems to enable high-impact science missions and the growth of a self-sustaining in-space economy. As a subsidiary of Ultra Safe Nuclear Corporation (USNC), USNC-Tech has leveraged the technologies and experience gains from the design, licensing, and deployment of terrestrial gas-cooled reactors for these space systems. The result is a family of reactor concepts and designs with shared technology bases able to meet near-term and long-term customer needs in the space arena. Key examples include the Pylon surface fission power reactor, the Ready 2 Demonstrate Technology On Orbit (R2DTOO) nuclear thermal propulsion reactor, and the chargeable atomic battery radioisotope heat source.

Mechanism analysis and experimental verification of electromagnetic sail, a new solar propulsion system without propellent

The electromagnetic sail, a new space propulsion system without propellant, which propels a spacecraft by deflecting solar wind with an electromagnetic field, is proposed. The thrust mechanism and the performance variations of the electromagnetic sail are analyzed with both particles in cell and magnetohydrodynamic models, and the interaction between the incoming plasma fluid and the electromagnetic field is investigated by tests on the ground experimental system. The simulation results demonstrate that a magnetosphere is created in front of the sail and its area increases with magnetic field intensity; with appropriate dipole moment, the thrust of the electromagnetic sail is higher than a pure magnetic sail or a pure electric sail. In test experiments, the plume structure of a helicon plasma source changes obviously with coil current and metal mesh voltage, which verifies the repulsion of the magnetic field and electric field to the incoming plasma. These results indicate that the combination of the magnetic sail and electric sail can increase the utilization efficiency of solar wind, which means the electromagnetic sail is a feasible propulsion system.

A drift kinetic model for the expander region of a magnetic mirror

We present a drift kinetic model for the free expansion of a thermal plasma out of a magnetic nozzle. This problem relates to plasma space propulsion systems, natural environments such as the solar wind, and end losses from the expander region of mirror magnetically confined fusion concepts such as the gas dynamic trap. The model incorporates trapped and passing orbit types encountered in the mirror expander geometry and maps to an upstream thermal distribution. This boundary condition and quasineutrality require the generation of an ambipolar potential drop of ∼5Te/e, forming a thermal barrier for the electrons. The model for the electron and ion velocity distributions and fluid moments is confirmed with data from a fully kinetic simulation. Finally, the model is extended to account for a population of fast sloshing ions arising from neutral beam heating within a magnetic mirror, again resulting in good agreement with a corresponding kinetic simulation.

Feasibility analysis of liquid propellant settling in the tanks of a space propulsion system after flying in zero gravity using a separating turn maneuver of the orbital unit

The paper presents results of preliminary design feasibility studies of using inertial settling of liquid propellant in tanks of space liquid rocket propulsion systems and separating from the propellant the ullage gases formed in zero gravity by means of centrifugal forces generated when the orbital unit performs the programmed turn maneuver proposed by the author. It shows that performing propellant settling and ullage gas separation by means of the separating turn maneuver makes it possible to significantly reduce propellant consumption in propulsion systems for attitude control and ullage engines. Key words: space liquid rocket propulsion system, main engine, separation of ullage gases from liquid propellant, separating turn maneuver of the orbital unit.

Sputtered highly effective iridium catalysts: a new approach for green satellite propulsion

This work demonstrated the large potential of sputtered iridium metal for catalytic reactions shown by the example of decomposition of hydrogen peroxide (H2O2) for space propulsion systems. For this purpose, iridium was coated onto Al2O3 pellets by a sputter process under varied process parameters. Depending on previously selected parameters, the obtained metal-loaded pellets offer closed- and/or open-shell structures. Catalytic productivity of these first-generation iridium-sputtered catalysts was estimated in laboratory experiments and compared to platinum-loaded pellets. Under optimized sputter-process conditions, the reactivity is significantly improved compared to the platinum-impregnated pellets. The better catalytic productivity can be explained by the increased active surface area of the iridium layers on the pellets. The surface morphology and the microstructure of the iridium coating can be actively controlled by the sputter pressure. The results are in accordance with the sputtering process pressure tendency described by the Thornton Structure–Zone Model.Graphical abstract

Acceleration characteristics of laser ablation Cu plasma in the electrostatic field

As a new concept of space propulsion system, laser-ablation propulsion has attracted more and more attention due to its characteristics of low power consumption, high specific impulse, variable and controllable thrust. With an aim to further raise up the movement velocity of plasma, we combine the laser with high-voltage electrostatic field to accelerate the Cu plasma induced by laser ablation. To demonstrate the acceleration characteristics of plasma under different electric field intensity, the plasma conductivity, plasma shockwave intensity and plasma plume movement process were tested using parallel electrode plate device, self-made torsion pendulum impulse test bench and high-speed ICCD camera. The results showed that the conductive current and impulse formed by the plasma obviously increased under the applied electric field. The images captured by high-speed ICCD camera showed the plasma cross-sectional area was 0.194 mm2 at 900 ns and 0.217 mm2 at 1600 ns when the electric field intensity was 0 V/mm. With the electric field intensity increased to 30 V/mm, the plasma cross-sectional area elevated to 0.280 mm2 at 900 ns and 0.288 mm2 at 1600 ns. The acquisitions prove that the idea of this paper is feasible and favorable, which provide a theoretical basis for the combination of laser ablation propulsion and electric field.