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Propeller is the crucial component of the fixed-wing aircraft. It provides thrust for fixed-wing aircrafts and has a great influence on the aerodynamic performance and carrying capacity of fixed-wing aircraft. For motor-driven micro-sized fixed-wing transport aircraft, the propeller needs to cooperate well with the motor to achieve high propulsion efficiency. In order to obtain the optimal design of propeller, experimental and numerical investigation was carried out and compared with three kinds of propeller. The tailor-designed propeller was proposed under specific conditions. Tailor-designed propeller better fits with the motor in the flight of the aircraft, improving the efficiency of the propellers, and increasing the carrying capacity of the aircraft. According to the characteristics and working conditions of the aircraft and the working characteristics of the motor obtained from the experiment, the aerodynamic shape of the propeller is initially designed based on the leaf theory and Betz conditions, then the parameters are continuously adjusted based on the numerical simulation results, and the results with the best performance are selected as the final result. Comparing the numerical simulation results and experimental results of tailor-made propellers and existing propellers under design conditions, the results show that, with the same power absorbed by the motor, the tailor-made propeller can provide greater thrust. The increase in thrust means that the aircraft can overcome greater resistance and generate greater lift, which improves the carrying capacity of aircraft. Numerical simulation and test flight methods were used to explore the impact of thrust improvement on aircraft lift and carrying capacity. [ABSTRACT FROM AUTHOR]

Wall pressure measurements are made in a convergent-divergent nozzle with strut injection to estimate the internal forces and moments. The strut facing the flow was of rectangular cross section and is placed at two thirds of the diverging length of the nozzle from the throat. The design exit Mach number of the nozzle is 1.84. The strut height varied at constant internal total pressure of 690 kPa. Experiments were conducted to explore the possibility of using a solid strut as an alternative to secondary fluid injection for thrust vector control. The wall pressure distribution of the nozzle is studied to interpret the flow interaction with the strut. The calculations based on the experimental data show that the presence of the strut and a variation in its height produce significant variations in side force and pitching moment which would be useful for thrust vector control. [ABSTRACT FROM AUTHOR]

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

We propose a horizontal-pendulum type thrust stand in which two-point acceleration measurement is applied to the null-balance method. The conventional displacement method was been applied to thrust measurement; thrust is determined by pendulum displacement which is caused by thrust. The null-balance method is also applied to the thrust measurement; a controller adjusts actuator force on the basis of pendulum displacement so as to null the pendulum displacement, and thrust is determined from actuator-driving current or voltage. However, both measurement methods can evaluate thrust variation up to one-third of the natural frequency that is determined by the configuration of the controller and pendulum. Hence, we proposed to apply acceleration measurement to the null-balance method, and showed that thrust variation could be measured up to 80 Hz. However, the proposed methods showed errors beyond 80 Hz; the error was caused by the translational flexibility of the torsional hinges. Then, we propose a new method that uses two-point acceleration measurement and adjustment of the center of gravity position to evaluate thrust variation beyond 80 Hz. The prototype evaluated thrust variation accurately beyond 200 Hz, although at 20-160 Hz, underestimating the amplitudes of reference thrust by 50 %. [ABSTRACT FROM AUTHOR]

The present research paper attempted to utilize a computational investigation for optimizing the fluidic injection angle effects on thrust vectoring. Simulation of a convergent divergent nozzle with shock-vector control method was performed, using URANS approach with Spalart–Allmaras turbulence model. The variable fluidic injection angle is investigated at different aerodynamic and geometric conditions. The current investigation demonstrated that injection angle is an essential parameter in fluidic thrust vectoring. Computational results indicate that optimizing injection angle would improve the thrust vectoring performance. Moreover, dynamic response of starting thrust vectoring would decrease by increasing nozzle pressure ratios and secondary to primary total pressure ratios. Also, shifting the location of fluidic injection towards the nozzle throat would have positive effect on response time. Additionally, the results of response time are more sensitive to primary and secondary total pressure ratios of nozzle and fluidic injection location than the fluidic injection angle. Furthermore, increasing fluidic thrust vectoring performance has negative impact on nozzle thrust at different expansion ratios. In addition, to guide the design and development of an efficient propulsion system, the current study attempted to initiate a database of optimum injection angles with different important parameter effects on thrust vectoring and nozzle thrust decline. [ABSTRACT FROM AUTHOR]

Abstract After billions of years of evolution, many creatures employing flapping wing in nature tend to have excellent flight capabilities. To understand the bionic wings flow mechanism will be helpful to design high performance underwater vehicles and new conception aircrafts. The geometric parameters, kinematic parameters and flow parameters have great effects on the bionic wings thrust performance. Facing the diverse parameters, it's very difficult to explore the three-dimensional (3D) bionic flapping wing flow mechanism with traditional numerical simulation method. In this paper, a general large-scale parallel solver using Immersed Boundary-Lattice Boltzmann Method (IB-LBM) was developed. The evolution procedures of the 3D flapping wing leading edge vortex and wake flow vortex structures were analyzed in detail. Our study explained the 3D flapping wing thrust performance variation with different wing shapes, aspect ratios and pitch-bias angles of attack. Using Chinese supercomputer TianHe-II presents a wide range of possibilities for the further development of parallel IB-LBM, employing tens of millions grids will help us to obtain more complete and accurate 3D flapping wing flow field information. It's indicated that the obtuse wing has the best thrust performance compared with other sharp wing shapes. With the increase of the aspect ratio, the thrust coefficient of flapping wing increases firstly and then decreases, and with pitch-bias angles of attack increases, the thrust coefficient decreases quickly or even shown resistance phenomenon at the large pitch-bias angel of attack. The discussion of these parameters will provide a theoretical basis for improving flapping-like vehicles propulsive performance. [ABSTRACT FROM AUTHOR]

Thrust zones in a deep-water fold-thrust belt, offshore northwestern Borneo, display prominent reflections that can be mapped through a three-dimensional seismic volume. Unlike faultplane reflections obtained from thrusts in other systems, these have positive polarity. Well data show that the reflections in the Borneo data set originate from fault-bound sandstone slices, with porosity-occluding calcite cement, entrained along the thrust zone. The thrust zone can support elevated fluid pressures beneath the fault with the cemented sandstones. Multiple sandstone slices indicate complex patterns of thrust zone localization, perhaps a common feature for deformation in sedimentary multilayers typical of many deep-water depositional successions. [ABSTRACT FROM AUTHOR]

In this paper, a fuel optimal rendezvous problem is tackled in the Hill-Clohessy-Wiltshire framework with several operational constraints as bounds on the thrust, non linear non convex and disjunctive operational constraints (on-off profile of the thrusters, minimum elapsed time between two consecutive firings...). An indirect method and a decomposition technique have already been combined in order to solve this kind of optimal control problem with such constraints. Due to a great number of parameters to tune, satisfactory results are hard to obtain and are sensitive to the initial condition. Assuming that no singular arc exists, it can be shown that the optimal control exhibits a bang-bang structure whose optimal switching times are to be found. Noticing that a system with a bang-bang control profile can be considered as two subsystems switching from one with control on to with control off, and vice-versa, a technique coming from the switching systems theory is used in order to optimise the switching times. [ABSTRACT FROM AUTHOR]

The electric propulsion system for CubeSat is required to achieve various missions. A miniature water ion thruster for CubeSats has been proposed and tested. To investigate the effect of the dissociation of water on its performance, a direct thrust measurement is required. In this study, a 100 μN-class thrust stand for 10 kg units was designed and integrated. It was calibrated using weights whose mass was measured in advance and achieved the measurement of the thrust of the miniature ion thruster using water. The range of measured thrust was 60-140 μN, and the average of its uncertainty was ± 7.6 %. The thrust coefficients were around 0.9, similar to or slightly less than that of xenon. This shows that the dissociation of water has little effect on the thrust. [ABSTRACT FROM AUTHOR]

The problem of the artificial frozen orbits around the Earth is investigated. To design such orbits with arbitrary orbital parameters, six control strategies are discussed taking into account the main zonal harmonics up to J4 and solar radiation pressure (SRP). The effect in the argument of periapsis due to the main zonal harmonics is revealed to be about three orders of magnitude greater than that of SRP. Two comparisons are given. To obtain longer lifespan of the spacecraft, continuous variable low-thrust control is implemented. Radial and/or transverse thrusts are used to achieve these orbits. Furthermore, explicit formulas for refinement of the control thrusts are given in order to overcome errors due to approximations arising from long and short periodic variations of a and e. Some concluding remarks can be inferred from the presented simulation; e.g. the energy consumed is saved efficiently when using the root mean squares rather than using averages, also when using coupled radial and transverse thrusts instead of using single radial or transverse thrust. To minimize the fuel consumption more efficiently, three new techniques are suggested. The calculations illustrated that AFOT-F Technique, where the transverse control thrust is constant, has the minimum control thrust required. [ABSTRACT FROM AUTHOR]

An annular lift fan aircraft configuration with high disk loading, low empty weight, and acceptable hover efficiency is assessed using numerical methods. The fans have short and very thin blades.With the fan disk loading increased from 28.37 to 91 lb/ft2, the figure of merit is 0.658 but the power loading drops from 9.03 to 4.65 lb/hp. However, when the aircraft is in ground effect within one aircraft radius to the ground plane, the power loading increases almost 50% to 6.77 lb/hp, which is comparable to helicopters and general aviation aircraft. The design features jet thrusters for forward flight propulsion, where the maximum requirement for jet thrust with the aid of a skirt in transition is 0.448 of the aircraft weight. Turbofan engines with a thrust-to-weight ratio of 0.3 and an augmentation ratio of 1.6-2.0 can satisfy this requirement and enable a top speed of Mach 0.62 (759 km/h) based on drag predictions. Shutters are proposed to close off the annular duct during cruise, and results show that they can close smoothly at the end of transition without causing oscillations of lift and drag. [ABSTRACT FROM AUTHOR]

Abstract Rotating detonation engine is a promising detonation-based propulsion system. In this paper, a series of numerical simulations of H 2 /air RDE are performed to show the plume flowfield of a RDE in detail and to obtain the pressure distribution on the end faces of the inner cylinder and the outer wall. The thrust force of the RDE is calculated with/without considering the two end faces to show its effect on the propulsive performance. Then the effects of the total massflow rate and the environment pressure on the propulsive performance are discussed. Under high environment pressure condition, the inner cylinder end face brings a notable disadvantage of the propulsive performance due to the strong suction effects. This disadvantage disappears under low environment pressure condition. The fluctuation of the total thrust force is dominated by the outer wall end face. As the massflow rate increases, the engine's thrust force becomes more steady. Under low environment pressure condition, the engine's thrust force is very stable and its fluctuation can be neglected. Under high environment pressure conditions, the use of the conventional thrust calculation formula overestimates the average thrust force and underestimates the unsteadiness of the thrust force obviously. [ABSTRACT FROM AUTHOR]

Aeroacoustic measurements and analysis have been made for an unshrouded rotor partially immersed in a planar equilibrium turbulent boundary layer at low Mach number. This configuration provides an idealized model of inflow distortion effects seen when a rotor is mounted adjacent to the hull or fuselage of a vehicle. At low and moderate thrust conditions, the rotor produces broadband noise organized into haystacks produced by large eddies of the ingested turbulence being cut multiple times by successive rotor blades. At high thrust, however, the acoustic signature changes and becomes louder and more tonal. This change is accompanied by separation of the boundary layer from the wall in the vicinity of the rotor blade disk. The separation region is highly unsteady and populated by intense vortex structures. Acoustic analysis suggests that blade–vortex interactions with these structures are the source of the additional tonal noise at high thrust. [ABSTRACT FROM AUTHOR]

A robotic fish is used to test the validity of a simplification made in the context of fish locomotion. With this artificial aquatic swimmer, we verify that the momentum equation results from a simple balance between a thrust and a drag that can be treated independently in the small amplitude regime. The thrust produced by the flexible robot is proportional to A2f2, where A and f are the respective tail-beat amplitude and oscillation frequency, irrespective of whether or not f coincides with the resonant frequency of the fish. The drag is proportional to U 0 2 , where U0 is the swimming velocity. These three physical quantities set the value of the Strouhal number in this regime. For larger amplitudes, we found that the drag coefficient is not constant but increases quadratically with the fin amplitude. As a consequence, the achieved locomotion velocity decreases, or the Strouhal number increases, as a function of the fin amplitude. [ABSTRACT FROM AUTHOR]

In this study, aerodynamic characteristics of a horizontal axis wind turbine (HAWT) were evaluated and discussed in terms of measured data in existing onshore wind farm. Five wind turbines (Tl, T2, T3, T4 and T5) were selected, and hub-height wind speed, Uq, wind turbine power output, P and turbine rotational speed, Q data measured from these turbines were used for evaluation. In order to obtain characteristics of axial flow induction factor, a, power coefficient, C;„ thrust force coefficient, CT, thrust force, T and tangential flow induction factor, a', Blade Element Momentum (BEM) theory was used. According to the results obtained, during a year, probability density of turbines at a rotational speed of 16.1 rpm was determined as approximately 45%. Optimum tip speed ratio was calculated to be 7.12 for most efficient wind turbine. Maximum Cp was found to be 30% corresponding to this tip speed ratio. [ABSTRACT FROM AUTHOR]

As one of the most important components of satellites, the thruster must maintain high reliability, but the assessment is difficult when lacking test failure data. When conducting hot test of the satellite thruster, we encountered the problem of reliability assessment on type-I censored data with only one failure, which is common for small sample tests. This paper proposes a novel interval statistic based reliability analysis method, which can fulfill the life information from the failure time to the censored time ignored by conventional methods, and improve the assessment accuracy. In this paper, a life distribution model is established by proper failure mechanism analysis and prior test information exploitation. As thrust chamber burning-through has been considered as the main failure mode, life test of the chamber coating was conducted under different thermal conditions to obtain the model parameter. Based on the interval statistic theory, a detailed derivation is illustrated. Then, reliability assessment and life prediction for the satellite thruster in both transfer orbit phase and synchronous orbit phase have been achieved, and the results show that our method performs very well, which provides an important way for dealing with test data with only one failure. [ABSTRACT FROM AUTHOR]

Dual-thrust rocket motors, with proper designs, can yield the desired boost-sustain thrust profiles for a variety of applications. Of the various designs that a dual-thrust rocket motor can have, the one with an intermediate nozzle provides a high boost-to-sustain thrust ratio as well as a stable operation. Depending on the internal conditions, the flow through the intermediate nozzle in the boost phase can be sonic, subsonic, or initially reversed. The present paper discusses the operation and performance of Dual-thrust rocket motors with subsonic intermediate nozzle. A mathematical model describing the internal ballistics of these dual-thrust rocket motors is derived from the basic governing equation. In addition, a set of experiments is conducted using a developed test motor to explore the physics of dualthrust rocket motors with subsonic intermediate nozzle and assess the mathematical model validity. Results of the theoretical and experimental investigations explain most of the internal ballistics characteristics that were not discussed before in the open literature. [ABSTRACT FROM AUTHOR]

In this paper, a novel analytical-based kinematic and dynamic modeling and control approach is developed for reference speed tracking of a multijoint robotic fish. The kinematic model is improved by the introduction of input signal for controlling body undulation amplitude without distorting the model structure. Next, the dynamic model is developed in the discrete-time domain primarily because fish swimming pattern generates oscillating continuous thrust where the effect of thrust over one cycle of undulation can be calculated through averaging. Hence, the dynamics can be discretized according to the undulation rate. The novelty in developing the discrete-time dynamical model is three fold. First, the thrust gradient is derived with assistance of Lighthill's large amplitude elongated body theory. More importantly, affine-in-control dynamical model form, essential for designing a real-time model-based control algorithm, is derived by inclusion of the updated kinematic model. Second, the thrust delay is studied via pulse input response of the motion system, and incorporated in the dynamical model. Third, drag force coefficient calculation is also proposed by comparing the thrust model and empirically recorded steady-state speeds of the robotic fish. Finally, discrete-time model-based sliding-mode control (SMC) is designed for speed tracking. Experimental results verify that SMC based on the proposed dynamical model, can significantly improve speed tracking performance of a multilink robotic fish. [ABSTRACT FROM PUBLISHER]

A class of abstract aerial robotic systems is introduced, the laterally bounded force vehicles, in which most of the control authority is expressed along a principal thrust direction, while along the lateral directions a (smaller and possibly null) force may be exploited to achieve full-pose tracking. This class approximates platforms endowed with noncollinear rotors that can modify the orientation of the total thrust in a body frame. If made possible by the force constraints, the proposed SE(3)-based control strategy achieves the independent tracking of position-plus-orientation trajectories. The method, which is proven using a Lyapunov technique, deals seamlessly with both underactuated and fully actuated platforms, and guarantees at least the position tracking in the case of an unfeasible full-pose reference trajectory. Several experimental tests are presented that clearly show the approach practicability and the sharp improvement with respect to state of the art. [ABSTRACT FROM AUTHOR]

In this paper, the nonlinear adaptive velocity and altitude tracking controller is developed for the longitudinal dynamics of generic air-breathing hypersonic flight vehicles. The proposed control scheme is designed using dynamic surface control method. The velocity and altitude subsystems are transformed into the linearly parameterized form for the convenience of adaptive law design. Both of the thrust and actuator constraints are explicitly considered. For thrust constraint, two cases are analyzed when the fuel-to-air ratio reaches its max and min values. A novel adaptive law is proposed to avoid over or less estimation in thrust saturation occasion. For actuator constraint, a magnitude and rate limiting filter is incorporated. The filter guarantees that the control signal is applicable for the actuator. It is shown that with the application of the proposed control scheme, all signals of the closed-loop system are uniformly ultimately bounded and the velocity and altitude tracking errors converge to a residual set which is arbitrarily small. Simulation results are demonstrated to show the effectiveness and superiority of the proposed control scheme. [ABSTRACT FROM AUTHOR]

With increased measurement tasks of space science, more requirements for the spacecraft environment have been put forward. Those tasks (e.g. the measurement of Earth's steady state gravity field anomalies) lead to the desire for developing drag-free control. Higher requirements for the thruster performance are made due to the demand for the drag-free control system and real-time compensation for non-conservative forces. Those requirements for the propulsion system include wide continuous throttling ability, high resolution, rapid response, low noise and so on. As a promising candidate, the cusped field thruster has features such as the high working stability, the low erosion rate, a long lifetime and the simple structure, so that it is chosen as the thruster to be discussed in this paper. Firstly, the performance of a new cusped field thruster is tested and analyzed. Then a drag-free control scheme based on the cusped field thruster is designed to evaluate the performance of this thruster. Subsequently, the effects of the thrust resolution, transient response time and thrust uncertainty on the controller are calculated respectively. Finally, the performance of closed-loop system is analyzed, and the simulation results verify the feasibility of applying cusped field thruster to drag-free flight in the space science measurement tasks. [ABSTRACT FROM AUTHOR]

DT-3 is a high performance low-freezing monopropellant developed in the 1980s in China, DT-3 is also the world's first practical use of low freezing point propellant. As a monopropellant for pose control, it is necessary to precisely control the thrust output. In this paper, a three-dimensional simulation is conducted, mass flow rate, preheating temperature and the size of the catalytic bed are studied for the controlling of thrust and start-up process. The results indicate that: When the mass flow is around 92.8 g/s, the thrust force reaches the designed thrust. Increasing mass flow, will lead to the temperature rise in the catalytic bed, and thus slightly accelerate the reaction process, eventually obtaining a relatively small propellant remaining at the catalytic bed exit position. Secondly, the increase in the preheat temperature facilitates faster entry of the thruster into the desired output state, while the impact on the final thrust is limited. Thirdly, the length of the posterior catalytic bed has an optimal value, the short catalytic bed causes the components to not react completely, excessive catalytic bed will affect the thrust output, and the best length is around 32 mm. [ABSTRACT FROM AUTHOR]

This paper analyzes the performance of an electric solar wind sail for generating and maintaining a heliocentric circular displaced orbit. Previous research on this subject was based on a simplified mathematical model of the spacecraft thrust. However, recent studies have proposed a more accurate algorithm for evaluating both the modulus and the direction of the propulsive thrust as a function of some important parameters related to the spacecraft attitude. Therefore, a reappraisal of the problem is motivated by the need to revise past results, taking into account new information available on the propulsion system. Within this context, this paper focuses on circular displaced orbits that are characterized in terms of orbital period, heliocentric distance and elevation angle. The attitude configuration and the value of the spacecraft characteristic acceleration required for orbital maintenance are calculated. An in-depth analysis of the linear stability of displaced orbits is given. It is shown that displaced orbits are unstable when the elevation angle exceeds about 20°. [ABSTRACT FROM AUTHOR]

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

Wind tunnel experiments are performed to investigate the aerodynamic characteristics of a quadrotor helicopter in a uniform flow. At first, the thrust and drag coefficients of a single rotor under low-Reynolds number conditions are measured to deduce the basic design parameters of the rotor. Second, the aerodynamic interference between two rotors in a tandem configuration is investigated. It is experimentally found that the thrust coefficient of the rear rotor is a maximum of 11% lower than that of the front rotor during forward flight. Moreover, it is found that the interference effect of the front rotor on the rear rotor can be predicted qualitatively using a theoretical formula on the basis of the Biot-Savart Law. Finally, it is shown that the theoretical predictions of the thrust coefficient and drag coefficient agree well with the experimental results of a quadrotor helicopter model within limited ranges of the inflow ratio and advance ratio. [ABSTRACT FROM AUTHOR]

Avian flight has an outstanding performance than the manmade flapping wing MAVs. Considering that the feather is light and strong, a new type of the flapping wing was designed and made, whose skeleton is carbon fiber rods and covered by goose feathers as the skin. Its aerodynamics is tested by experiments and can be compared with conventional artificial flapping wings made of carbon fiber rods as the skeleton and polyester membrane as the skin. The results showed that the feathered wing could generate more lift than the membrane wing in the same flapping kinematics because the feathered wing can have slots between feathers in an upstroke process, which can mainly reduce the negative lift. At the same time, the power consumption also decreased significantly, due to the decrease in the fluctuating range of the periodic lift curve, which reduced the offset consumption of lift. At the same time, the thrusts generated by the feather wing and the membrane wing are similar with each other, which increases with the increase of flapping frequency. In general, the aerodynamic performances of the feather wing are superior to that of the membrane wings. [ABSTRACT FROM AUTHOR]

In this paper, an efficient aerodynamic analysis framework is developed for the conceptual design of lift-fan type aircraft. The lift-fan type aircraft includes and uses a rotor in the body or wing for vertical takeoff and landing, and flies differently from a rotorcraft in cruise/forward flight. Since the lift-fan type aircraft can fly like a fixed wing vehicle during cruise past the transition flight phase, it is possible to overcome the range and velocity limits of conventional rotorcraft. To investigate a wide range of design options during the conceptual design, rapid analyses of various design candidates are significant as high-fidelity analysis may be too expensive. A cost-efficient, lift-fan aircraft aerodynamic analysis (LFAAA) method is developed which consists of a linear potential flow solver with viscous correction, the Blade Element Momentum Theory (BEMT), a Power Induced Effects (PIE) module and momentum drag correction. The greatest advantage of the LFAAA method is low computational cost with considerable solution accuracy as long as the flow conditions are in relatively linear regions, by which the individual component analysis method is derived. Wind-tunnel data of various geometries including an isolated propeller of a UAV, Lockheed Martin innovative control effector (ICE), and YAV-8B Harrier II aircraft are used to validate each component of the LFAAA method, respectively. For the validation of the LFAAA method for comprehensive analysis, a complete geometry of NASA's 1/8 scale model of fan-in-wing (FIW) aircraft is solved and the computation results are validated with the existing wind-tunnel data and unsteady RANS computations. Results predicted by the LFAAA method show good agreements with high-fidelity URANS and wind-tunnel data over the linear flow regimes where no flow separation is observed. [ABSTRACT FROM AUTHOR]

In this paper, the combustion process of liquid propellant ammonium dinitramide (ADN)-methanol aqueous solution in a small space thruster is investigated experimentally and numerically. The phase Doppler anemometer measurement is used to investigate the injection process and the results are used as the liquid boundary condition in the simulation to minimize uncertainties. Hot fire tests have also been operated to estimate the performance of the thruster. To simulate the complicated processes in the thruster, a modified multi-component droplet model is employed to simulate the evaporation process considering the interaction between the liquid droplet and the porous media. A non-equilibrium model for the porous media is used to describe the heat transfer in the catalyst bed. A reduced chemical mechanism containing 18 species and 40 reactions is used to simulate the reactions of gas-phase ADN and methanol. The results show that the decomposition of ADN and the oxidization of methanol do not happen synchronously in the chamber. The decomposition of ADN takes place near the inlet while methanol is oxidized downstream of the porous media. The establishment of pressure in the thruster depends on the evaporation and reaction processes and the temperature changes relatively slowly because the porous media increases the thermal inertia of the whole system. [ABSTRACT FROM AUTHOR]

A control parameter direct optimization method for the optimal short-range elliptic orbit rendezvous problem using on–off constant thrust is proposed. This method is based on analytical state propagation expressions for linear relative motion under on–off constant control acceleration. This avoids numerical integration for the orbit dynamic equations. Normalized optimization variables are introduced to replace switching times. Then this problem is transformed into a nonlinear programming problem with bound constraints on the optimization variables and terminal equality constraints. The along-track and cross-track thrusts are used for the in-plane and out-of-plane motions, respectively. Both the minimum-fuel and minimum-time solutions are obtained by using the Matlab function fmincon. Numerical examples are provided to verify the proposed method. The results indicate that the proposed method is accurate and fast. [ABSTRACT FROM AUTHOR]

Three different types of high power Hall thrusters--anode layer type, magnetic layer type with high specific impulse, and magnetic layer type with dual mode operation (high thrust mode and high specific impulse mode)--have been developed, and the thrust performance of each thruster has been evaluated. The thrust of the anode layer type thruster is in the range of 19-219 mN, with power in the range of 325-4500 W. The thrust of the high specific impulse magnetic layer type thruster was 102 mN, with specific impulse of 3300 s. The thrust of the bimodal operation magnetic layer thruster was 385mN with specific impulse of 1200 s, and 300mN with specific impulse of 2330 s. The performance of these thrusters demonstrates that the Japanese electric propulsion community has the capability to develop a thruster for commercial use. [ABSTRACT FROM AUTHOR]

Geological interpretation and seismic imaging on data acquired from complex media, especially when accompanied by a thick layer of salt and/or anhydrate, is always a controversial task. The aim of this study is to introduce a method for seismic imaging of complex structures and study the folding scenario of the Zagros overthrust in SW Iran based on the improved seismic image processed by the proposed method. A diffraction based version of the common reflection surface stack method, called the common diffraction surface (CDS), was selected to enhance quality of seismic image from the Zagros fold-thrust belt. The 2D CDS method was also developed here to 3D partial CDS. Consequently, three sets of 2D and a set of 3D seismic data were processed with the conventional Kirchhoff prestack time migration and the partial common diffraction surface methods. Geological models released in previous studies of the Zagros overthrust were used as the base model of interpretation. Structural interpretation on seismic sections with new method was in accordance with previous geological models and improved in some parts where the new method could enhance quality of the seismic image. Seismic structural interpretation shows the inner part of the belt was subjected to folding, uplift and erosion. This mechanism constituted a resistant mass in the Dezful Embayment in front of the SW moving thrust and folds. This resistant mass formed tightly folded or thrusted and highly shortened structures in the Izeh zone and High Zagros area. In the oil rich zone of the Dezful Embayment, the thick evaporate-rich Gachsaran Formation, decouples the folding above it from the folding below, making seismic imaging of potential traps beneath the detachment horizons extremely difficult. [ABSTRACT FROM AUTHOR]

Based on the inherent coupling of system states, backstepping controllers are designed for underactuated spacecraft hovering without either the radial or in-track thrust and in the presence of input saturation and unmatched disturbances. Feasible hovering positions for either saturated underactuated case are derived. A Lyapunov-based approach is used to prove the asymptotical convergence of state errors and system robustness against unmatched disturbances. The validity of theoretical analysis is verified by simulations in a J2 -perturbed environment. [ABSTRACT FROM AUTHOR]

We have developed a novel continuation technique to solve optimal bang-bang control for low-thrust orbital transfers considering the first-order necessary optimality conditions derived from Lawden's primer vector theory. Continuation on the thrust amplitude is mainly described in this paper. Firstly, a finite-thrust transfer with an “On-Off-On” thrusting sequence is modeled using a two-impulse transfer as initial solution, and then the thrust amplitude is decreased gradually to find an optimal solution with minimum thrust. Secondly, the thrust amplitude is continued from its minimum value to positive infinity to find the optimal bang-bang control, and a thrust switching principle is employed to determine the control structure by monitoring the variation of the switching function. In the continuation process, a bifurcation of bang-bang control is revealed and the concept of critical thrust is proposed to illustrate this phenomenon. The same thrust switching principle is also applicable to the continuation on other parameters, such as transfer time, orbital phase angle, etc. By this continuation technique, fuel-optimal orbital transfers with variable mission parameters can be found via an automated algorithm, and there is no need to provide an initial guess for the costate variables. Moreover, continuation is implemented in the solution space of bang-bang control that is either optimal or non-optimal, which shows that a desired solution of bang-bang control is obtained via continuation on a single parameter starting from an existing solution of bang-bang control. Finally, numerical examples are presented to demonstrate the effectiveness of the proposed continuation technique. Specifically, this continuation technique provides an approach to find multiple solutions satisfying the first-order necessary optimality conditions to the same orbital transfer problem, and a continuation strategy is presented as a preliminary approach for solving the bang-bang control of many-revolution orbital transfers. [ABSTRACT FROM AUTHOR]

A linear voice coil motor (VCM) with forward winding for linear vibro-impact systems is proposed in this paper. Generally, only increasing the air-gap flux density of VCM, which ignores its bad thrust characteristic condition, is limited to obtain a higher thrust density, due to its intermittent duty and reciprocating linear motion. Thus, the analysis on the thrust characteristic is of momentous significance for VCM design. Considering this fact, we try to adjust the thrust characteristic to obtain the maximum thrust density. Magnetic-circuit analysis method is utilized to analyze the thrust characteristic along with a given total working stroke. Besides, finite-element analysis is also used in the selection of parameters. A series of prototype VCMs for experiments is made. The results from both theoretical simulations and practical experiments show increase of thrust density and decrease of thrust ripple compared with the original design, and prove the superiority of the proposed structure. [ABSTRACT FROM PUBLISHER]

Our project is to design a multirole fighter jet aircraft. A multirole fighter jet aircraft is an aircraft designed to perform different roles in warfare. The air-toair combat role has been normally performed by fighter aircraft. In addition, a multirole fighter jet aircraft has secondary roles such as air-to-surface attack, intercepting, reconnaissance, escorting and air cover. The term multirole has been reserved for aircraft designed with the aim of using a common airframe for multiple tasks where the same basic airframe is adapted to a number of differing roles. The main motivation for developing multirole aircraft is cost reduction in using a common airframe. The project report comprises of a literature survey of about 10 fighter aircrafts. After this an appropriate aerofoil is selected and its important parameters are computed. Then, performance graphs are drawn and a 3-view diagram of the aircraft is drawn. Finally, model is made and tested. [ABSTRACT FROM AUTHOR]

This study investigated an approach to reduce the number of actuators used for internal pressure control and thruster allocation in a multinozzle solid propulsion system. In the proposed design, the throat areas of four divert nozzles are controlled by only three actuators, and chamber pressure maintenance and thrust distribution are achieved by controlling the throat areas. Using the proposed actuator set, thrust allocation can be accomplished in a more efficient way than when independent actuators are employed for each nozzle. [ABSTRACT FROM AUTHOR]

Purpose The purpose of this paper is to present a novel high-precision relative navigation method for tight formation-keeping based on thrust on-line identification.Design/methodology/approach Considering that thrust acceleration cannot be measured directly, an on-line identification method of thrust acceleration is explored via the estimated acceleration of major space perturbation and the inter-satellite relative states obtained from space-borne acceleration sensors; then, an effective identification model is designed to reconstruct thrust acceleration. Based on the identified thrust acceleration, relative orbit dynamics for tight formation-keeping is established. Further, using global positioning system (GPS) measurement information, a modified extended Kalman filter (EKF) is suggested to obtain the inter-satellite relative position and relative velocity.Findings Compared with the normal EKF and the adaptive robust EKF, the proposed modified EKF has better estimation accuracy in radial and along-track directions because of accurate compensation of thrust acceleration. Meanwhile, high-precision relative navigation results depend on high-precision acceleration sensors. Finally, simulation studies on a chief-deputy formation flying control system are performed to verify the effectiveness and superiority of the proposed relative navigation algorithm.Social implications This paper provides a reference in solving the problem of high-precision relative navigation in tight formation-keeping application.Originality/value This paper proposes a novel on-line identification method for thrust acceleration and shows that thrust identification-based modified EKF is more efficient in relative navigation for tight formation-keeping. [ABSTRACT FROM AUTHOR]

A high aspect-ratio foil is known to be advantageous in terms of both thrust and efficiency in flapping propulsion. However, many species of fish have evolved a low aspect-ratio hydrofoil, which naturally leads one to search for its physical advantages in locomotion. Here we study the flow physics of a hydrofoil in angular reciprocating motion with negligible free-stream velocity to reveal the effects of an aspect ratio on hydrodynamic performance. By establishing a scaling law for the thrust of a foil of general shapes and corroborating it experimentally, we find that the thrust of an angularly reciprocating foil is maximized at a low aspect ratio of 0.7 while hydromechanical efficiency continuously increases with an aspect ratio. This result suggests that a low aspect-ratio foil can improve thrust produced by the foil when they start from rest, but at the expense of efficiency. [ABSTRACT FROM AUTHOR]

The minimum-time constant-thrust circular orbit rephasing problem is studied using a curvilinear relative motion description for dynamics. The resulting optimal control problem in the thrust orientation is formulated using the indirect method and investigated both analytically and numerically. By linearizing the relative motion equations, a key nondimensional parameter is identified, which characterizes the duration of the maneuver and its qualitative structure when changing front a short-maneuver to a long-maneuver time regime, passing through a transition zone. Approximate analytical solutions are obtained for the short- and long-maneuver regimes, providing an accurate estimation of the minimum maneuver time and a clear understanding of the evolution of the optimal thrust profile, which approaches two opposite bang-bang structures in the limit cases. The nonlinear problem is investigated numerically and shown to be fully characterized by two nondimensional parameters instead of one. Finally, a comparison w ith different two-impulse maneuvers is conducted. The results can be used for constructing good initial guesses in more complex optimization problems and to support preliminary mission design. [ABSTRACT FROM AUTHOR]