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As a novel micro-propulsion system for small satellites (from micro to nano), the ionic liquid electrospray propulsion system is a promising candidate. However, performance measurement and evaluation of the Ionic Liquid Electrospray Thruster (ILET) is one of the most challenging issues for practical application, due to the difficulties in the development of a prototype and direct measurements of micro-thrust and small flow rate. To address this issue, a Modular Ionic Liquid Electrospray Thruster (MILET) prototype is constructed, and a diagnostic system for thrust and mass flow rate is specially developed based on an analytical balance method. With the ionic liquid 1-ethyl-3-methylimidazolium tetrafluoroborate as the propellant, a series of experiments is carried out on the MILET prototype under a wide operating condition through changing the applied voltage to control the thrust. Under different applied voltages, the thrust and the mass flow rate of the propellant are directly measured. The propulsive performance parameters of the thruster, such as thrust, specific impulse, thrust-to-power ratio, thruster efficiency, etc., are comprehensively analyzed. Then, a performance comparison is made between the MILET and other representative ILETs. With a relatively low applied voltage ranging from 1550 V to 2000 V, the MILET achieves a quasi-constant specific impulse of 1263 s with the averaged thrust-to-power ratio of 65.2 μN/W and thruster efficiency of 40.7%. The performance of ILET is also compared with other typical electric propulsions. The results demonstrate that the ILET exhibits an excellent ability of minimalization with high specific impulse and thruster efficiency, which guarantees a great superiority in micro propulsions. Finally, the ways to further improve the performance of ILET are discussed, which further confirms the potential prospect of ILET. The present result helps to advance the development and application of ILET.

The amalgamation of metallic alloys and fiber-reinforced composites has helped Fiber Metal Laminates (FMLs) like Carbon fiber reinforced aluminum laminates (CARALL) and Glass laminate aluminum reinforced epoxy (GLARE) overcome the limitations of standalone metals and composites. As a result, they have found increasing applications in the aircraft and defense industries. Moreover, the inhomogeneous nature and poor machinability of the materials make hole drilling a challenging task. Therefore, to enhance drilling performance, the paper reports a systematic analysis on the effect of spindle speed and axial feed on key performance measures, including thrust force, machining temperature, surface roughness, hole size, burr size, and chip morphology. Experimental outcomes indicate a significant reduction in the thrust force while utilizing higher spindle speed (4000 rev/min) and lower axial feed (0.1 mm/rev), while the surface finish improved under high feed conditions (0.4 mm/rev). The analysis revealed that machining temperature increased with the employment of higher spindle speed and lower axial feed. Higher spindle speeds and axial feeds are desirable from the perspective of hole accuracy as they help produce holes within H9 diameter tolerance. Burr size was verified to be larger at the hole exit compared to hole entry, with the size of the burr increasing with an increase in spindle speed and axial feed. The results show that the axial feed was the significant variable affecting chip size followed by spindle speed. Higher axial feed (0.4 mm/rev) and spindle speed (4000 rev/min) helped improve the chip breakability, thus helping in better chip evacuation.

High-Altitude Long-Endurance (HALE) solar-powered Unmanned Aircraft Vehicles (UAVs) can utilize solar energy as power source and maintain extremely long cruise endurance, which has attracted extensive attentions from researchers. Trajectory optimization is a promising way to achieve superior flight time because of the finite solar energy absorbed in a day. In this work, a method of trajectory optimization and guidance for HALE solar-powered aircraft based on a Reinforcement Learning (RL) framework is introduced. According to flight and environment information, a neural network controller outputs commands of thrust, attack angle, and bank angle to realize an autonomous flight based on energy maximization. The validity of the proposed method was evaluated in a 5-km radius area in simulation, and results have shown that after one day-night cycle, the battery energy of the RL-controller was improved by 31% and 17% compared with those of a steady-state strategy with a constant speed and a constant altitude and a kind of state-machine strategy, respectively. In addition, results of an uninterrupted flight test have shown that the endurance of the RL controller was longer than those of the control cases.

The article presents a new approach for optimizing the trajectories of spacecraft with finite thrust. The recently proposed angular variable – the auxiliary longitude – is considered as an independent variable. The advantages of using this variable to optimize a trajectory with a fixed angular distance and free transfer duration are shown. A complete system of necessary optimality conditions for the considered representation of the optimal control problem is presented. An approach to solving the minimum-fuel problem is considered that is based on the sequential solution of the problems of optimizing the limited power trajectories and the trajectories with a constant exhaust velocity. This approach includes verifying the existence of a solution to the minimum-fuel problem by solving the thrust minimization problem. We present a method for optimization of the finite-thrust trajectories with constant exhaust velocity, based on the application of the maximum principle and the continuation method, which does not require the user to specify an initial guess for the initial values of costate variables. The article presents numerical examples to illustrate the application of such approach, and discusses the properties of the obtained optimal trajectories.

A comparative evaluation of conventional drilling and helical milling to process holes in AZ31 Magnesium alloy is reported. The results highlight the advantages of helical hole milling over the mechanical drilling process in terms of lower machining forces. Besides, the process is successful in reducing the machining temperature and size of the burrs formed, in addition to improved surface finish. Further, analysis of the helical hole milling process demonstrated that the axial pitch has a dominating influence on the thrust force, whereas the feed value greatly influences the radial force. The spindle speed is found to influence the machining temperature significantly. The chip morphology is found to be dependent on feed and axial pitch, with the feed value influencing the chip segmentation and breakability. The absence of burn marks on the chip surface indicates a lower risk of fire hazard.

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.

A novel turbofan Direct Thrust Control (DTC) architecture based on Linear Parameter-Varying (LPV) approach for a two-spool turbofan engine thrust control is proposed in this paper. Instead of transforming thrust command to shaft speed command and pressure ratio command, the thrust will be directly controlled by an optimal controller with two control variables. LPV model of the engine is established for the designing of thrust estimator and controller. A robust LPV H∞ filter is introduced to estimate the unmeasurable thrust according to measurable engine states. The thrust estimation error system is proved to be Affinely Quadratically Stable (AQS) in the whole parameter box with a prescribed H∞ performance index γ. Due to the existence of overdetermined equations, the solving of controller parameters is a multi-solution problem. Therefore, Particle Swarm Optimization (PSO) algorithm is used to optimize the controller parameters to obtain satisfactory control performance based on the engine’s LPV model. Numerical simulations show that the thrust estimator can acquire smooth and accurate estimating results when sensor noise exists. The optimal controller can receive desired control performance both in steady and transition control tasks within the engine working states above the idle, verifying the effectiveness of the proposed DTC architecture’s application in thrust direct control problem.

Rock-soil interface mixed ground (RSI) is often encountered in tunnel construction. The excavation loads of tunnel boring machines (TBMs) are controlled by the interaction characteristics between TBM and rock/soil layers. The different properties of rock and soil cause the varying interaction range and stress distribution. Currently, there have been several studies available to estimate excavation loads under RSI, and the conclusion is that the total loads increase with increasing the rock layer proportion in the excavation face. However, the previous studies cannot take the difference of rock/soil properties into account, except for the calculation of cutters loads. Therefore, the interaction characteristics between RSI and TBM is unclear. This paper analyzes the interaction characteristics between TBM’s main components and complex geological conditions (e.g., layered soil, layered rock, and RSI condition). A model is proposed to calculate the total thrust and total torque assuming quasi-static equilibrium of the tunneling equipment. The rationality and applicability of the model are discussed and verified by two typical projects. Furthermore, the geological adaptability is discussed in terms of the excavation difficulty and the matching relationship between total torque and total thrust. The results indicate that when the rock layer proportion in the excavation face increases, the reduction of overall extrusion and friction loads is 1.5 times higher than the increase of disc cutters breaking load. The total loads and the ratio of the total torque to total thrust decrease approximately linearly. There is a power function relationship between the excavation difficulty index and the penetration depth. The results of this study provide an important reference for the total loads design of equipment propulsion systems and the parameter adjustment during tunnel construction.

In recent years, Chinese Long March (LM) launchers have experienced several launch failures, most of which occurred in their propulsion systems, and this paper studies Autonomous Mission Reconstruction (AMRC) technology to alleviate losses due to these failures. The status of the techniques related to AMRC, including trajectory and mission planning, guidance methods, and fault tolerant technologies, are reviewed, and their features are compared, which reflect the challenges faced by AMRC technology. After a brief introduction about the failure modes of engines that can occur during flight, and the fundamentals of trajectory planning and joint optimization of the target orbit and flight path, an AMRC algorithm is proposed for geostationary transfer orbit launch missions. The algorithm evaluates the residual performance onboard, and plans new objectives and corresponding flight path by iterative guidance mode or segmented state triggered optimization methods in real-time. Three failure scenarios that have occurred during previous LM missions are simulated to check the robustness of the algorithm: imminent risk of the boosters’ engines, thrust drop during the first stage of flight, and being unable to start the engine during the second stage. The payloads would fall from space according to the current design under these conditions, but they were saved with the AMRC algorithm in the simulations, which allowed the rocket to get into the target orbit as intended or the payloads were deployed in other orbits without crashing. Although spaceflight can be very unforgiving, the AMRC algorithm has the potential to avoid the total loss of a launch mission when faced with these kinds of typical failures.

Accurate prediction of tip vortices is crucial for predicting the hovering performance of a helicopter rotor. A new high-order scheme (we call it WENO-K) proposed by our research group is employed to minimize numerical dissipation and extended to numerical simulation of unsteady compressible viscous flows dominated by tip vortices over hovering rotors. WENO-K is referred to as an adaptively optimized WENO scheme with Gauss-Kriging reconstruction, and its advantage is to reduce dissipation in smooth regions of flow while preserving high-resolution around discontinuities. Here WENO-K scheme is adopted to reconstruct left and right state values within the Roe Riemann solver updating the inviscid fluxes on a structured dynamic overset grid. To minimize the accuracy loss for high-order reconstruction on artificial boundaries of overset grid, a method of multilayer fringes is proposed to carry out interpolation between background grid and blade grid. Massively parallel computing considering automatic load balance on averagely partitioned overset grid is developed to reduce the wall-clock time of an unsteady simulation. Numerical results for Caradonna-Tung (C-T) rotor in hover at the conditions of subsonic and transonic tip Mach numbers show that the thrust coefficient error for the result of WENO-K scheme is no more than 3%. Compared with WENO-JS scheme, WENO-K scheme achieves about 40% improvement on accuracy of predicting rotor thrust with only 4.1% extra computational cost. More importantly, WENO-K scheme can capture more sophisticated unsteady flow structures and resolve tip vortices to a larger wake age with an increment of about 270° compared to WENO-JS scheme.

The process of starting the rotating detonation engine initially filled with air in low outer pressure conditions was studied. Such engines can be used for bringing satellites into the Earth orbit. A three-dimensional numerical simulation of an engine combustion chamber with a rotating detonation wave of a cylindrical type with an internal body fed by a hydrogen-oxygen mixture or an acetylene-oxygen mixture was performed. The influence of the lengths of the combustion chamber channel on the stability of the detonation wave and thrust characteristics was considered. An author's computer code was used in simulations, the code was based on a model of multicomponent gas dynamics with chemical transformations and turbulence. The code was tested by comparing it with experiments and analytical solutions for simple cases. To describe the kinetics of acetylene combustion, a simplified chemical scheme of interaction of 10 components was used, without “heavy” radical species: [C2H2, CO, CO2, H2, O2, H2O, OH, O, H, N2]. The influence of the chosen kinetic mechanism in case of hydrogen-oxygen mixture on the simulation results was studied. Various modes of the process in the detonation engine were obtained, including those with a stable detonation wave.

Nitrous oxide (N2O) is a green propellant with excellent application prospects. A sub-Newton N2O monopropellant thruster with inner-heater and a N2O self-pressurization stable supply system with regenerative heat compensation are designed in this paper. The experimental research of the thruster is described, including measurements of preheating power, activation temperature, vacuum thrust, specific impulse, life-span and pulsed operation performance. By inserting the heater into the catalyst-bed, preheating efficiency of the heater is significantly improved compared to the thruster with outer-heater. Thus, the preheating power demand of the thruster is successfully reduced to around 10 W. The mean vacuum thrust of 322 mN is attained and the corresponding specific impulse reaches 162 s at the mass flow rate of 0.2 g/s. Successful activation temperature of 523 K is achieved, and the activation performance of the thruster is affected by the loading factor. A long term hot-firing test longer than 12000 s is attained. The pulsed operation performance of the inner-preheating thruster is also studied by measuring chamber pressure. Impulses with different magnitudes are produced by adjusting the opening duration of the control valve. A minimum impulse of 81 mN∙s is attained. Finally, the performance of the thruster is evaluated by comparison with other thrusters of the same type. The results indicate that the proposed thruster with inner-heater is superior in terms of preheating power, activation temperature and specific impulse performance.

The largest population of space debris is observed in the low-Earth-orbit region with altitudes of 600–1000 km and inclinations of 80–100°. As follows from the predictive calculations, space debris removal from this region exactly is of the first priority. For this reason, our article discusses the problems of controlling the motion of service spacecraft – space debris object cluster when the space debris object is removed from a low orbit with high inclination. In the conceptual design of the spacecraft presented hereinafter, the electric propulsion system contains two electric propulsion thrusters of stationary plasma thruster type, each of which is mounted with the possibility of independent rotation respect to two mutually perpendicular directions. An ion beam injector forms an ion beam with 65 mN of thrust. Spacecraft motion control, including the thrust vector control, is provided by coordinated rotation of thrusters. For such concept, a strategy and an algorithm for controlling the motion of the spacecraft center of mass are proposed, which are based on the control of the spacecraft lateral and longitudinal motion that is understood as the control of the relative distance between the spacecraft and the space debris object. The strategy of spacecraft lateral motion control assumes that due to the thrusters' rotation the spacecraft should be displaced in such a way that the vector of the relative range between the spacecraft and the space debris object coincides with the direction of the transversal of the orbital coordinate system. For the practical implementation of the control, an algorithm has been developed for controlling the thrusters' rotation to create the required values of the spacecraft acceleration projections. The mathematical model includes the equations of motion of the spacecraft center of mass taking into account the aerodynamic deceleration. The solar panels' orientation to the Sun is provided by the spacecraft roll rotation relative to its longitudinal axis and turning the solar panels relative to their axis of rotation. The change in the ion beam momentum transfer is considered using a simple simulation model. Mathematical modeling was performed for the following parameters of the initial orbit: altitude: 800 km; eccentricity: 0.0005; and inclination: 900. For the proposed control algorithm, a range of possible thrust of a single thruster was 51–60 mN. The simulation results show that the proposed control can be used to remove space debris object from low orbits with high inclination, taking into account the roll spacecraft rotation to keep the solar panels’ orientation to the Sun.

With the development of the aircraft gas turbine engine, a control system should be able to achieve effective thrust control to gain better operability. The main contribution of this paper is to develop a novel direct thrust control approach based on an improved model predictive control method through a strategy that reduces the dimension of control sequence. It can not only achieve normal direct thrust control tasks but also maximize the thrust level within the safe operation boundaries. Only the action of switching the objective functions is required to achieve the switch of these two thrust control modes while there is no modification to the control structure. Besides, a shorter control sequence is defined for multivariable control by updating only one control variable at every simulation time instant. Therefore, the time requirement for the solving process of the optimal control sequence is reduced. The proposed controller is implemented to a twin-spool engine. Simulations are conducted in the wide flight envelope, and results show that the average time-consumption can be reduced up to 65% in comparison with the standard model predictive control, and the thrust can be increased significantly when maximum thrust mode is implemented by using engine limit margins.

The alternative working modes and flexible working states are the outstanding features of an adaptive cycle engine, with a proper control schedule design being the only way to exploit the performance of such an engine. However, unreasonable design in the control schedule causes not only performance deterioration but also serious aerodynamic stability problems. Thus, in this work, a hybrid optimization method that automatically chooses the working modes and identifies the optimal and smooth control schedules is proposed, by combining the differential evolution algorithm and the Latin hypercube sampling method. The control schedule architecture does not only optimize the engine steady-state performance under different working modes but also solves the control-schedule discontinuity problem, especially during mode transition. The optimal control schedules are continuous and almost monotonic, and hence are strongly suitable for a control system, and are designed for two different working conditions, i.e., supersonic and subsonic throttling, which proves that the proposed hybrid method applies to various working conditions. The evaluation demonstrates that the proposed control method optimizes the engine performance, the surge margin of the compression components, and the range of the thrust during throttling.

A high thrust liquid oxygen-kerosene fuelled engine is being developed for future launch vehicle applications. Comprehensive hot testing of engine at sea level conditions is required for qualification and acceptance of flight engine and launch vehicle stage. The engine and stage will be tested in a specialized rocket engine and stage hot test facility which deflects the flame safely to avoid any hazard to engine and adjoining mechanical structures. The facility is equipped with water jets to sufficiently cool the rocket plume before impinging on deflector pit walls. The high temperature and high-pressure rocket exhaust jet is cooled with sufficient quantity of water to maintain the impingement wall temperature within safe working limits. In this study, a multi-phase computational fluid dynamics (CFD) methodology is developed for design of water-cooling jet configuration of deflector facility. Multi-phase model validation study using full phase treatment and droplet assumption for water injection is performed prior to water jet optimization analysis of high thrust LOX-kerosene test facility. Validation studies based on simulation and analysis of existing test stand is developed. A discrete particle (DPM) based two phase methodology is chosen to model water injection in RANS framework. The optimum cooling configuration is determined iteratively through numerous multi-phase simulations. The multi-phase analysis showed ineffectiveness of initial water injection scheme which results in facility wall temperature above safe limit. A multi-plane water injection scheme is devised there on with provision of water injection at two axial locations from the nozzle exit. A detailed parametric multi-phase analysis is carried out for design optimization of multi-plane water injection scheme and to determine thermal condition of the test facility. Analysis showed effective rocket plume cooling and lower facility wall temperature with two plane water injection schemes in comparison to initial single plane design configuration. The study highlights the role of CFD based methodology to formulate an engineering solution for design optimization of LOX-kerosene engine hot test facility.

Nowadays, the development of low Earth orbits’ (LEO) - up to 2000 km - small satellites (SSs) is of particular interest. Such SSs are supposed to use for effective telecommunication systems and advanced Earth research missions. When developing space missions of LEO SSs, designers and operators encounter a number of the following technical problems: an inability to use a single propulsion system for all flight operations, such as transfer from parking orbit to the target one, maneuvers between orbits, altitude control, orbit maintenance, attitude control, de-orbiting; a limited lifetime of conventional (Hall-Effect and Ion type) Electric Propulsion Systems (EPs) in low Earth orbits; high energy consumption of existing EPs; unloading of SSs' attitude control systems; de-orbiting of the SSs from the target orbit at the end of the active life to reduce contamination of LEO. The multidirectional electrodeless plasma thrusters could be the solution to overcome these problems. For example, such propulsion systems (PSs) can generate thrust in two directions. Using three bi-directional plasma thrusters in one PS, it is possible to create a plasma propulsion module with six thrust vectors that allows performing all of the SSs' space mission tasks where propulsion systems are needed while having a long-term service life in the upper atmosphere, low energy consumption – from 100 W, - and compact size – from 1U - which allows using this module on board of the SSs. This paper describes the first prototype of the six-directional plasma propulsion module. The focus of this paper is the study of the bi-directional wave plasma source with an open-ended gas discharge chamber and magnetic nozzles for plasma flow control – plasma flows acceleration and locking. The example of this source demonstrates the possibility of creating two thrust vectors using the single bi-directional wave plasma source and allows suggesting the creation of the six-directional wave plasma propulsion module for SSs.

With the discovery of 1I ’Oumuamua and 2I Borisov, interstellar objects (ISOs) have been thrust into the forefront of planetary science, and the robotic exploration of ISOs is now of significant interest. With current propulsion technology, rendezvous with these objects is likely infeasible, and thus the maximum science return results from a combination of flyby spacecraft and impactor spacecraft, as demonstrated by the Deep Impact encounter with comet Tempel 1. Interstellar objects by nature have hyperbolic orbits and pass through the inner solar system at extreme velocities, which presents several unique challenges to navigation. This paper first establishes the possible encounter space intercepts with ISOs may present. We then evaluate the ability for intercepting spacecraft to navigate to ISOs and autonomously track/impact targets across the possible encounter space, depending on key navigation hardware such as camera and attitude determination system.

In 1969, John Ford Fishback proposed a solution for the interstellar ramjet with magnetic scooping. With the exception of two papers published in the early 1970s, this model has not been revisited up to now for viability. Here, we analyse Fishback’s axial scoop field configuration. Computer simulations prove its functionality, albeit with absurdly long magnetic coils when reasonable thrust is to be achieved. Based on the Maxwell stress tensor we obtain a much lower upper bound for the cut-off speed at a given acceleration of the ramjet than published before. Our results indicate that there are severe engineering problems with the magnetic collection process. It is very unlikely that even Kardashev Civilizations of type II might build magnetic ramjets with axial solenoids.

The detonation wave in a rotating detonation engine is highly adaptable to the incoming flow, making the wave easier to control. In this study, a numerical simulation method is used to analyze the working process and flow field structure of a rotating detonation model engine with dual cavity injection of an H2/air mixture by controlling the injection pressure ratio of the dual cavity and the number of detonation wave heads. It is found that the rotating detonation engine offers the possibility to control the thrust vector with two different modes. The first is a one-cycle alternate control mode with a small injection pressure ratio. Here two deflections occur in different directions occur across one detonation wave propagation cycle, but the overall deflection direction is in the low-pressure region. The second is a one-way control mode, with a large injection pressure ratio, and the deflection direction towards the low-pressure region. For the multi wave-mode, it belongs to one-way control mode because of constant deflection direction in the low-pressure area. From the perspective of thrust distribution along the circumference, the one-way control strategy satisfies the ability of a rotating detonation thrust vector control.

In order to find a possible relation between the reservoir level and the deformation of the Sanalona dam, the horizontal and vertical displacements for a period of 3.5 years using GPS observations were calculated. The GAMIT/GLOBK scientific software was used for the relative processing of GPS observations and annual speeds over the dam of −6.77 mm/yr, −9.95 mm/yr and −1.62 mm/yr were calculated for the N, E and U components, respectively. In addition, the GNSS reflectometry inversion technique was applied to determine the daily reservoir level of the dam, the values obtained were compared with the real reference values of graduated rulers in situ. The GNSS-R achieved a correlation coefficient ( ρ ) of 0.97 with a root-mean-square error (RMSE) of 1.36 m and a regression slope of 0.932 m/m. The deformation in the x Dam component obtained a greater relation with the reservoir level due to the thrust force generated on the crest. The results demonstrate that it is possible to use a GPS antenna to carry out two monitoring processes together on a dam: (1) determination of displacements and (2) calculation of reservoir level.

Medium-frequency (mid-frequency) vibration analysis of complex structures plays a prominent role in automotive, aerospace, mechanical, and civil engineering. Flexible beam structures displayed by the traditional Euler–Bernoulli pillar hypothesis have been broadly utilized in different designing issues. A kinematic speculation made in the Euler–Bernoulli shaft hypothesis is that the plane areas of a pillar typical to its impartial pivot remain planes after the bar encounters twisting disfigurement, which ignores shear deformity. The expanded DTFM decides the recurrence reaction of a bar structure in a careful and insightful structure, in any recurrence locale covering low, center, or high frequencies. In the mean time, the proposed technique gives the nearby data of a bar structure, like uprooting, shear misshapening, bowing second and shear power at any area, which in any case would be extremely challenging with energy-based strategies. amplifiers by and large accept that a confined main thrust can spread energy uniformly across the outside of a board. Notwithstanding, examinations have shown that board vibrations stay confined around the driving point at high frequencies, Energy spreading will possibly happen when the board is incited in a recurrence district with a low thickness of modes, as numerous modes impelled together will join to frame a band-restricted delta work at the area of the main impetus. A quantitative proportion of limitation is presented, in light of the proportion of the energy contained in a little locale around the driving point to the energy contained in the whole board. Analyses show that the limitation transcends a basic recurrence, and this basic recurrence connects well with forecasts of the great recurrence plate district utilizing standard modular cross-over measurements.

Air Foil Thrust Bearings (AFTB) are self-acting hydrodynamic members that support axial loads of high-speed rotating turbo-machineries. The load bearing capacity of such bearings mainly depends on the foil structure. The proposed Foil Thrust Bearing of 60 mm in diameter, designed and fabricated for performance evaluation, consists of a single top foil with an inner edge height equal to that of outer edge. This paper focuses on optimizing the foil stiffness in terms of foil thickness for steel and copper foil material with 60° and 90° sector angles. Various configurations of Gas Foil Thrust Bearings (GFTB) are tested for structural stiffness and thrust loads at higher operating speeds with an instrumented test rig dedicated for the purpose. Both geometric and operating parameters are varied to determine the best possible foil stiffness that enhances the load carrying capability of the Air Foil Thrust Bearing operating under hydrodynamic condition. CFD Simulation is also carried out to indicate the pressure distribution on the foil surface for the said bearing at actual boundary conditions.

In the present study, wind tunnel experiments were performed to investigate the power and thrust characteristics of a wind turbine model subjected to a variety of surge and sway motions. The turbine model was mounted on a translational stage to simulate the surge and sway motions at prescribed amplitudes and frequencies. The power output and rotor thrust were measured in each sets of the motion frequencies and amplitudes, respectively; and the measurements were also performed in a base-fixed turbine for comparison. Results show that the surge and sway motions with various amplitudes and frequencies all have slight impact on the mean power output and mean rotor thrust of a wind turbine. However, the rotor thrust fluctuations of the surging turbine, not only appeared significantly higher than those of the swaying turbine and base-fixed turbine at the same sets of amplitudes and frequencies, but also increased with the surge frequency and surge amplitude. Moreover, for both the surging and swaying turbine, the rotor thrust fluctuated with the same period as the surge and sway motions, respectively; and the power spectral density of the rotor thrust both exhibited a peak at the motion frequency and its odd multiple frequencies.

Future asteroid landing and sample return missions will seek for landing sites with high scientific value, such as hazardous terrains, which means that the spacecraft needs the ability of autonomous hazard avoidance. Nowadays, the guidance algorithm based on artificial potential function plays an increasingly important role in hazard avoidance, but the local minimum problem makes that the spacecraft cannot reach the desired target landing point in some complex terrains. In this paper, a novel hazard avoidance guidance method which improves the traditional artificial potential function is developed. This method focuses on the impact zone analysis of hazards and defines anti-collision zone. In addition, a new repulsive potential function in logarithmic form is designed with continuity and fast numerical change rate in the anti-collision zone on this basis, and the problem of spacecraft falling into local minimum zone in complex terrains can be effectively avoided. Considering the practical application of constant thrust engine, the hazard avoidance control law with constant thrust sliding mode is designed to reduce switching frequency and fuel consumption. The performance of the proposed guidance law is verified through a set of simulations, as well as the global stability of the control system under uncertainty and perturbation conditions.

This study investigates the drilling operation on Inconel 718 superalloy by using cryogenic treated drill bit. Thrust force, torque and surface roughness are analyzed to understand the effects of the cryogenic condition during CNC drilling operation. Taguchi’s methodology has been employed for planning of experiments. Drill bits are conditioned under two different environment i.e. cryogenically treated with single tempered and cryogenically treated with double tempered. The process parameters are optimized by using Whale optimization algorithm for obtaining minimized surface roughness along with maximization of torque and thrust force during drilling operation. The ANOVA results show that feed rate is the most contributing significant factor for thrust for its maximization while spindle speed is the most contributing significant factor for torque for its maximization. Also, tool condition is the most contributing significant factor for surface roughness for its minimization. Lastly, confirmatory test were conducted to find out the significance of the predicted values obtained using the optimization methods.

In this paper, the primary emphasis was on the modelling and study of aircraft winglets. Drag mitigation is a major problem in aerodynamic engineering. A gadget called a winglet is mounted vertically at a series of angles on the end of the aircraft wing to minimise drag. By decreasing the aircraft drag, the winglet configuration would decrease the fuel consumption and making the aircraft more stable during takeoff, as well as giving the aircraft engine a longer life by reducing the load on the engine thrust. The aim is to design and simulate a using CATIA V5, which is used to build the various winglet models (blended winglet, rated winglet) at different winglet angles (350, 650 and 900) with angle of attack and ANSYS is used to measure and simulate the winglet model. In this study it was observed that wings with winglets produce higher lift force/Drag for the CFD and its static FE analysis study. Its effectiveness is reduced by a certain degree of angle of attack and by rising further to a higher angle of attack. A potential solution for enhancing the aerodynamic performance of an aircraft seems to be the examined idea of variable angle winglets.

The solid fuel ducted Ramjet systems (SFDR) are extremely capable of propelling high- speed missiles. They have a specific application in the configuration of air-to-air missiles with a high degree of interception. It is a modification from the traditional solid fuel ramjet system by using an external duct for thrust variation. This system provides higher effectiveness against maneuvering jets compared to single and dual pulse rocket system as it can vary thrust by varying fuel and air quantity for combustion. Even though with high effectiveness, an enemy fighter jet that is having higher thrust engines can outmaneuver SFDR by gaining significant lift during the end-stage which reduces its killing probability; this emphasized the development of a new missile system that provides more lift during the end-stage. A new SFDR concept is proposed in the present work that can gain more lift. The provision of a deflector is made in the duct to study the lift characteristics. Optimum lift force was obtained from the detailed analysis. Moreover, the study was extended by changing the deflector dimensions. Fluent fluid flow analysis is performed in simulation software named ANSYS to study the lift characteristics. Before that, 3D modelling is done with the help of NX software.

The widespread use of unmanned aerial vehicles (UAVs) has gained significant importance given their low cost and range of possible applications. Nevertheless, design methodologies integrating aerodynamic and structural evaluations for UAV propellers have not been widely explored. A proper design of UAV propellers provides compelling energy savings, and the selection of structurally adequate blades ensures UAV integrity. This work presents a methodology for the aerodynamic design of propellers using parametric and high-fidelity tools coupled with a structural evaluation scheme using FEM. For this aim, different propeller configurations were determined using BEMT and MIL approaches. Furthermore, a baseline airframe and usual operating conditions for monitoring tasks were selected and evaluated throughout several optimized propeller configurations. The design obtained through the two-dimensional aforesaid approach is evaluated using numerical simulation, coupling the aerodynamic loads with the structural design through FSI in ANSYS 18. Subsequent aerodynamic evaluation revealed congruence between BEMT and CFD results, with a 9% maximum percentage error between results. Moreover, the findings of the studied case scenarios showed an improvement in thrust generation capabilities for a 1.48 m diameter 3 bladed propeller, with an 80% efficiency. Results for the structural assessment indicated acceptable stress concentrations for all configurations at the design point and an overall better structural performance for a 0.99 m diameter 3 bladed propeller. To summarize, the main contribution of this work is a design methodology tailored for UAV propellers able to predict their performance and optimal configurations using a fast and accurate enough calculation scheme for preliminary design stage.

The transfer between libration orbits in the Mars–Phobos system is challenging under the strong gravitational perturbation of Mars and Phobos. In this paper, the outer branches of manifolds of libration orbits around Phobos are employed for impulse and continuous thrust transfer between orbits, respectively. The 1:1 resonant orbits are found to be appropriate intermediate orbits for impulse transfer, which can connect the manifolds of various kinds of libration orbits. For the continuous thrust transfer, the convex optimization method is adopted to design the low-thrust trajectory in the three-body problem. The Poincare section is established to find the best matching points on the outer manifolds. The L 1 / L 2 homoclinic and L 1 − L 2 heteroclinic transfer trajectories are designed and discussed. The simulations show the feasibility of the transfer by outer manifolds with both impulse and continuous thrust. The proposed method can avoid the potential hazard caused by the complicated dynamics near Phobos and provide a reliable transfer method, which can be used in future Phobos exploration missions.

Hovering is an effective and challenging way for asteroid exploration. In this paper, the adaptive quasi-periodic switching control method for body-fixed hovering of asteroids with low fuel consumption and strong robustness using constant thrust is proposed. By analyzing the active control process and the drift process of the spacecraft in the phase plane, constant thrust control is applied in the appropriate orientation to maintain the desired hovering state above the asteroids. Considering the uncertainty of the dynamic environments near the asteroids, the apex position of drift process in the phase plane is predicted, meanwhile, the compensation conditions and the calculation formula for the constant-thrust engine ignition are given. The simulation results show that the quasi-periodic switching control method can reduce the fuel consumption of the spacecraft compared with other methods, moreover, the adaptive quasi-periodic switching controller can effectively improve the robustness of the system.

The gas-particle two-phase flow in a convergent-divergent nozzle is encountered in many applications like solid-propellant rocket motors (SRMs). The “Particle-free Zone” adjacent to the nozzle wall is crucial for the nozzle performance which is closely related to the specific impulse of the motor. Though it has been usually observed by many investigators, the existence of the particle-free zone in different nozzles and its effect on nozzle performance are still ambiguous. In this research, we investigated the two-phase flow with monodispersed micron particles in a nozzle using a two-way coupled Eulerian-Lagrangian model and found that the particle-free zone does not exist for smallest particles (dp = 1.0 μm). Moreover, we built a theoretical model for predicting the extent of the particle-free zone by calculating the trajectory of the particle closest to the nozzle wall. The effects of the particle size, total pressure, total temperature and nozzle geometry on the particle-free zone were studied. Finally, we analyzed the nozzle performance and found that the gas velocity at the outlet and the thrust do not change with particle size monotonously, and the medium-sized particles cause the largest thrust loss.

Multi wind tUrbine Simulation Tool (MUST) is a new CENER's in-house code, based on NREL's OpenFAST code, designed to allow the aero-hydro-servo-elastic modelling of novel multi-rotor wind turbine concepts. MUST also includes a new aerodynamic module, called AeroVIEW (Aerodynamic Vortex fIlamEnt Wake) based on an implementation of the Free Vortex filament Method (FVM) combined with an unsteady Lifting Line (LL). This model allows to represent the influence existing between all the rotors, in each one of them. In this article, a validation of AeroVIEW performance for single and multi-rotor configurations is presented. First, a study of the wake properties (wake geometry, core radius and circulation of tip vortex filament) obtained with AeroVIEW has been successfully compared with several MexNext III cases. Then, an analysis of the rotor wakes interaction, in different arrays of wind turbines, has been made, obtaining power and thrust increases, due to this rotor-wake interaction, for different conditions. The increases obtained in power are very similar to results of high fidelity tools found in the literature. Moreover, this beneficial effect on the power generated, has a counterpart in the average thrust, whose increase is around half of the power production increase.

In this study, the changing performance characteristics of a solid-fuel scramjet with a size-fixed aft channel for a long-duration operation were investigated using numerical methods. The results show that a solid-fuel scramjet with such a chamber configuration can continuously provide effective thrust over a long working period, followed by a long tail section of thrust. The effective working period lasts from the beginning of the operation to the moment when the bottom wall covered by the solid fuel is exposed to hot gas. The fluctuation in the thrust is significant due to the relatively large fluctuation in the equivalence ratio. The solid fuel utilization is relatively low; approximately 49% of the solid fuel remains after the effective working period. The configuration used in this study limits the mass flow rate passing through the front part of the cavity significantly, thereby forcing the core of the main recirculation zone to move downward rapidly, with undesirable effects on the equivalence ratio fluctuation and utilization of the front part of the solid fuel.

Control technologies are innovated to satisfy increasingly complicated control demands of gas turbine engines. In terms of limit protection control, a novel model-based multivariable limit protection control method, which is achieved by adaptive command reconstruction and multiple-control loop selection and switch logic, is proposed in this paper to address the problem of balancing smaller thrust loss and safe operations by comparing with widely-used Min-Max logic. Five different combination modes of control loops, which represent the online control loop of last time instant and that of current time instant, is analyzed. Different command reconstructions are designed for these modes, which is based on static gain conversion of amplitude beyond limits by using an onboard model. The double-prediction based control loop selection and switch logic is developed to choose a control loop appropriately by comparing converted amplitude beyond limits regardless of one or more parameters tending to exceed limits. The proposed method is implemented in a twin-spool turbofan engine to achieve limit protection with direct thrust control, and the loss of thrust is improved by about 30% in comparison with the loss of thrust caused by Min-Max logic when limit protection control is activated, which demonstrates the effectiveness of the proposed method.

The poor quality of the micro-hole in the conventional drilling (CD) of 3D needled Cf/SiC composites seriously limits its application in modern industry. Ultrasonic assisted drilling (UAD) is a hybrid machining technology that combines CD with ultrasonic vibration. When processing holes in the Cf/SiC composites, the UAD has great advantages of improving hole wall quality and reducing cutting force, and it has been widely applied to Cf/SiC manufacturing. This study mainly focuses on the influence of amplitude in UAD on the manufacturing of micro holes in the Cf/SiC. By analyzing the micro-morphology of the hole wall and outlet edge under different amplitude in UAD, the influence of amplitude on the fiber removal mode and material damage mechanism is studied. Meanwhile, the thrust force in CD and UAD is also obtained. The results show that the defects of the hole are reduced with the alteration of the fiber removal mode in UAD. And with the increase of the amplitude, the hole quality is further improved. However, the excessive amplitude leads to the damage of the matrix. The thrust force curve shows a “multi-peak” shape with a regular interval due to the layered structure of Cf/SiC in CD and UAD. Compared with the CD, the thrust force of UAD is reduced by up to 70.1% as the amplitude increases from 3 μm to 9 μm. Furthermore, the hammering effect caused by ultrasonic vibration alleviates the material damage around the outlet edge in UAD. With the increase of amplitude, the matrix damage and the fracture of the fiber around the outlet are alleviated. The fracture feature of the fiber around the outlet is more regular than the CD. Consequently, UAD is a beneficial method for Cf/SiC composites micro-drilling with better hole quality and it plays an important role in the future of Cf/SiC processing.

In this paper, the kerosene/air rotating detonation engines(RDE) are numerically investigated, and the emphasis is laid on the effects of total pressures and equivalence ratios on the operation characteristics of RDE including the initiation, instabilities, and propulsive performance. A hybrid MPI + OpenMP parallel computing model is applied and it is proved to be able to obtain a more effective parallel performance on high performance computing(HPC) systems. A series of cases with the total pressure of 1 MPa, 1.5 MPa, 2 MPa, and the equivalence ratio of 0.9, 1, 1.4 are simulated. On one hand, the total pressure shows a significant impact on the instabilities of rotating detonation waves. The instability phenomenon is observed in cases with low total pressure (1 MPa) and weakened with the increase of the total pressure. The total pressure has a small impact on the detonation wave velocity and the specific impulse. On the other hand, the equivalence ratio shows a negligible influence on the instabilities, while it affects the ignition process and accounts for the detonation velocity deficit. It is more difficult to initiate rotating detonation waves directly in the lean fuel operation condition. Little difference was observed in the thrust with different equivalence ratios of 0.9, 1, and 1.4. The highest specific impulse was obtained in the lean fuel cases, which is around 2700 s. The findings could provide insights into the understanding of the operation characteristics of kerosene/air RDE.

The partial space elevator is a space transportation system that conveys cargo by a climber moving along the tether that connects the main satellite and the end body. This work proposed a novel prescribed performance based dual-loop control strategy for configuration keeping of the partial space elevator in cargo transportation. The control loop I controls the libration angle of the climber by a newly proposed prescribed performance control law to track the desired state with bounded performance. The Lyapunov stability of the control law is proved analytically. Control loop II stabilizes the overall configuration of the PSE by applying thrust and deploying/retrieving the tether simultaneously at the end body. To enforce the constraint that the total tether length is unchanged at the end of transportation, an event-triggered loaning control algorithm is proposed to further improve the Control loop II. Numerical simulations validate the prescribed performance based dual-loop control strategy. The results show that the stable configuration of the PSE is kept effectively with no libration motion in the period of cargo transportation.

Electric thrusters are considered to be the most promising propulsion system of choice for space gravitational wave detection. However, the requirement of high accuracy and drag-free control results in significant challenges in thrust measurements. To better understand the thrust noise characteristics, this work proposes a time-frequency-domain method for thrust noise analysis, i.e., a scalogram method attained through wavelet analysis. Unlike the conventional Fourier transform method, it can provide 3-D relationships for the time-frequency and thrust noise. A free-state thrust measurement experiment identified that the environmental temperature produces the dominant impact in the thrust noise. During a 70,000 s measurement, even without thruster operation, the thrust noise increased from 0.3 μN/Hz1/2 to 19 μN/Hz1/2 as the environmental temperature increased from −1 °C to 3 °C. The corresponding frequency of peak thrust noise for the entire period tends to be the same of 0.1 Hz. Additionally, a Hall thruster thrust measurement experiment was conducted to verify the method's time-varying analysis capability for thrust noise. The complete time-frequency thrust noise variation throughout the entire operating process of a Hall thruster was observed. The results at different stages of thruster operation indicate that thrust noise has the greatest influence 150–300 s after ignition. This is because the thrust stand will take hundreds of seconds to reach thermal equilibrium after ignition. Another finding during the Hall thruster operation is that the 0.1 Hz frequency corresponding to the peak thrust noise is highly consistent with the frequency of peak thrust noise in the free-state thrust experiment. This reveals the fact that the 0.1 Hz frequency should be the dominant frequency for thrust noise caused by temperature drift in our thrust measurement system.

A comprehensive benchmarking study has been carried out to determine the influence of the mesh formulation and chip separation methods on the reliability and accuracy of finite element modelling of large diameter drilling operations. The Coupled Eulerian–Lagrangian and the updated-Lagrangian (with element deletion) formulations available in ABAQUS/Explicit, together with the updated-Lagrangian (with re-meshing) formulation in DEFORM 3D are compared by simulating through-coolant drilling of AISI 1045. The Johnson–Cook damage model was implemented by a sub-routine in DEFORM 3D to ensure a consistent damage model is implemented across the formulations. Experimentally measured drilling thrust force, torque, and chip thickness values were used to compare the models performance and assess the accuracy of the predictions. The updated-Lagrangian with dynamic re-meshing was found to be the best performing methodology concerning the accuracy of predictions, whilst the Coupled Eulerian–Lagrangian methodology significantly under-predicted the drilling thrust force and torque. Due to numerical instabilities and computational cost, the updated-Lagrangian with element deletion method is not recommended to model large diameter drilling.

This paper proposes an improved method for the preliminary evaluation of minimum-propellant trajectories to Near-Earth Asteroids (NEAs). The method applies to missions from Earth to asteroids with small eccentricity and inclination. A planar and a plane-change problem can be distinguished. In the planar problem, the solution assumes that multiple burn arcs are performed in correspondence of the apsides of the target asteroid in order to change the initial spacecraft orbit (i.e., Earth’s orbit) into the target one. The number of arcs is established once the time of flight is given (1 burn at each apsis per revolution, 1 revolution per year can be assumed). The length and propellant consumption of each arc to attain the required changes of semi-major axis and eccentricity are computed by a procedure based on Edelbaum’s approximation, which is well-suited to the problem at hand, as eccentricity changes are expected to be small for feasible missions. No numerical integration is required, but only the numerical solution of a three-unknown algebraic system is needed, making the procedure extremely fast. Plane change is taken into account assuming a constant out-of-plane thrust angle during each burn. A previous simple formulation used an averaged thrust effect over one revolution and neglected the fact that plane changes are more effective at the nodes. Several improvements are here introduced, which greatly increase the method accuracy. The influences of the eccentricity change, the angle between the asteroid line of nodes and line of apsides, and the expected length of the arc are considered: In fact, when the eccentricity is small, the thrust arc can be performed at the nodes where the inclination is efficiently changed, with little penalty in the planar maneuver. An efficient plane change is also performed when the angle between the asteroid line of nodes and line of apsides is small and/or the length of the arc is large, because, in this case, the node is comprised in the apsidal burn. A simple corrective formula accounts for this effect. The new method shows remarkable accuracy. The results comparison with solutions obtained with an indirect optimization method for a set of more than 60 NEAs shows a 0.95 correlation coefficient in the propellant masses. The estimation error is below 10% for 75% of the targets, below 15% for 95% of the targets, and always below 20%.

The propellantless working principle of a solar sail requires the total flight time to be minimized when looking for the optimal trajectory to reach a given target state. In this work the solar sail steering law is found by applying a Q-law algorithm, which aims at driving the spacecraft toward the final (target) orbit by decreasing the distance between actual and desired states or by increasing the rate of change of the state variables. A formulation of the Q-law algorithm for a solar sail-based mission is given, which accounts for the constraints along the transfer trajectory imposed by the sail thrust model. The performance of the proposed procedure, which represents the novelty of this work, is checked in some potential solar sail mission scenarios, including coplanar interplanetary transfers and the exploration of outer Solar System regions.

In the present work, the aerodynamic design of a 10 kW horizontal axis wind turbine rotor is performed using both ideal and actual rotor theories. The obtained nonlinear blade profile is optimized in order to enhance the aerodynamic performance and to ease the fabricating complexity. A unique linearization approach is employed to linearize the chord and twist distribution by dividing the congruent line of both ideal and actual models into equal divisions. The points along the identical tangent line are considered as floating new blade roots, whereas the blade tip was kept fixed based on the primary design. The linear profile based on the new value of blade root is described using algebraic equations. The local element torque, capacity factor, and the annual energy production based on the Weibull distribution are adopted to determine and assess the optimal blade profile. Both CFD investigation and the FEA analysis are performed in order to evaluate both primary and optimized blades. The aerodynamic and aeroelastic comparison in terms of power output, thrust force, blade tip deflection, and the equivalent stress distribution of both blade profiles has been done over a wide range of incoming wind speeds. Results show an enhancement in the aerodynamic performance in terms of power coefficient up to (5.9%) compared to the primary blade design. Moreover, the optimized blade has shown less tip deflection by 27.92% than the primary blade at low wind speed.

In this study, friction drilling was performed for AA7075 - T6 alloy under dry, minimum quantity lubrication (MQL), heat, and heat + MQL process environments. The aim is to achieve the optimum bushing profile by increasing the productivity of friction drilling in different process environments. Friction drilling experiments were carried out at constant speed and feed rate; the effects of the drilling environment on thrust force, temperature, hole bushing profile geometric properties, hole surface quality, thread stripping strength, and micro hardness values were investigated. The test results revealed that the drilling environment directly affects the formation of heat and thrust forces during the process. The best surface roughness (Ra) value was achieved in the heat environment with 2.61 µm. Thread stripping experiments showed that the heating process improved the thread stripping strength of AA7075-T6 by 14.65% when compared with the dry environment. The results revealed that the main damage mechanism occurred in the transition zone from the heat-affected zone (HAZ) to the base material (BM) in the tensile strength test on the heat process environment specimen, while the roots of threads were affected in other process environments. The different microhardness values were measured in the thermo-mechanically affected zone (TMAZ), HAZ, and BM formed during the process.

Due to the errors in calculating external load and thrust system parameters, traditional dynamic positioning model of oil drilling platform has poor positioning accuracy, making it difficult to meet the demand of oil exploitation. Therefore, a new dynamic positioning model of offshore oil drilling platform is designed based on OIPSO algorithm. Based on the mathematical model of offshore oil drilling platform, the external load (wind, wave and ocean current) is calculated, and the movement of the platform is simulated, thus the motion equation of the platform is constructed. Based on this, OIPSO algorithm is used to distribute the thrust force for dynamic positioning, and to determine the value range of decision variables. Then combined with PID controller, the movement of the offshore oil drilling platform can be controlled, so as to realize the dynamic positioning of the platform. The experimental results show that compared with the traditional model, the dynamic positioning error of the proposed model is smaller, and the model has higher validity and feasibility.

Aiming to maximize the aerodynamic performance of the Distributed Electric Propulsion (DEP) aircraft, a hybrid design framework which focuses on the aerodynamic performance of the propeller/wing integration has been developed and validated numerically. Variable-fidelity modelling for propeller aerodynamics has been used to achieve computational efficiency with reasonable accuracy. By optimizing the aerodynamic loading distributions on the tractor propeller disk, the induced slipstream is redistributed into a form that is beneficial for the wing downstream, based on which the propeller blade geometry is generated through a rapid inversed design procedure. As compared with the Minimum Induced Loss (MIL) propeller at a specified thrust level, significant improvements of both the lift-to-drag ratio of the wing and the propeller/wing integrated aerodynamic efficiency is achieved, which shows great promise to deliver aerodynamic benefits for the wing within the propeller slipstream without any additional devices.

Longitudinal torsional coupled rotary ultrasonic assisted drilling (LTC-RUAD) technology is introduced to improve the surface roughness of the hole wall and solve the tear, burr and delamination of carbon fiber reinforced polymers (CFRPs) induced by large thrust force and torque during conventional drilling (CD) in this study. First, the defect suppression mechanism of CFRPs hole was proposed via the analysis of the types and forming mechanism of CFRPs hole defects using a tapered drill-reamer (TDR) in CD and the characteristics of LTC-RUAD. Then, the corresponding drilling experimental platform was built based on the novel designed LTC-RUAD vibration actuator, while drilling experiments involving T700S-12K/YP-H26 CFRPs specimens with different process parameters were performed by adopting the different ultrasonic vibration amplitude (UVA) in the longitudinal and torsional direction. The thrust force and torque, quality of the hole wall and holed defects obtained in both CD and LTC-RUAD were compared at the same process parameter. The experimental results show that compared with CD, the maximum average thrust force and torque can be reduced by approximately 30% and 40%, respectively. The surface roughness, defects at the exit of the drilled hole show a “concave” trend with the increase of UVA, and the burr factor, tear factor and delamination factor are reduced by 62.56%, 76.6%, 69.67% when the longitudinal UVA is approximately 7–9 μm, respectively. The forming mechanism of hole defects in CD and the defect suppression mechanism of CFRPs hole in LTC-RUAD are verified to be correct.

An atmosphere-breathing electric propulsion system (ABEP) uses the rarefied atmospheric molecules as the propellant for the electric thruster. In the best case, it can allow spacecraft (S/C) complete a long-time mission in very low Earth orbit (VLEO) without carrying any propellant. However, previous studies have shown that the thruster lifetime is limited by electrode erosion due to the aggressive gases, such as the atomic oxygen in VLEO. This paper design and analyse an inductively coupled plasma generator (ICPG) for ABEP system, and it is electrodeless to avoid electrode erosion. Based on the COMSOL Multiphysics, the simulation model of ICPG is established and analyzed. Multiple physical fields are coupled in this model including plasma, magnetic field, and flow field. The simulation model is verified by the experimental results of the literature. Driving by the radio frequency power, the plasma power distribution and thrust estimation of the ICPG under different input conditions (different drive powers and different operation gases) can be obtained. The results show that the ICPG may have a better performance under the condition of low mass flow and high drive power to compensate the drag. The gas species and their proportions have also great influence on the behaviours of ICPG. This paper can provide a convenient and efficient method to study the ICPG products for ABEP.

The rocket exhaust, as a source of the disturbance, provides a way to examine physical and chemical phenomena in the ionosphere. Since the disturbance morphological characteristics are closely related to rocket orbit, the ionospheric perturbation produced by the rocket exhaust considering the trajectory effect is simulated. The three-stage booster rocket trajectory is calculated, considering multi-stage thrust, aerodynamic drag, gravity and apparent force. Based on trajectory calculation and the electron density depletion theory of neutral gas injected into ionosphere, the exhaust is divided into a series of mass elements released at different time and different positions along the trajectory. The theory considers the charge exchange between the ambient O+ ions and the high-speed exhaust molecules (primarily H2O) in the exhaust plume. The simulation results show that the depletion structure restricted by the trajectory are dramatically different from the structure produced by a point source release, meaning that it no longer appears as an isotropic quasi-symmetric distribution, but the overall orientation of the depletion structure is consistent with the trajectory direction. Furthermore, the ray tracing technique is used to simulate the propagation of radio waves passing through the depletion structures. Due to the steepness of the depletion boundary and the difference in the shape of the boundary, the “focusing” effect of the rays which though the disturbance structure restricted by the trajectory disappear. These characteristics results can be an important reference for rocket detection and identification.

Analytical synthesis of extremal transfer trajectories with up to three intermediate thrust arcs between elliptical orbits is presented using Mayer’s variational problem. Solution uniqueness requires three such arcs connected via intermediate transfer orbits. Junction conditions provide state vector continuity at junction points. Fuel efficiency is compared with impulsive trajectories.