Tracking control problems for space robots are studied under conditions without speed feedback signals. An adaptive RBF neural network control method with a speed observer is proposed. Specially, we conduct the following. 1) A dynamic model of space robots is established. 2) A speed observer based on a neural network is designed to reconstruct speed information. 3) A controller based on a neural network is designed to compensate the nonlinear model of system. 4) A weight adaptive learning laws of the neural network is designed to ensure on-line tuning without an off-line learning phase. 5) The uniformly ultimately bounded state of the closed-loop system is proved based on Lyapunov theory. Simulation results show that the adaptive neural network control method with the speed observer can achieve good precision. This has important engineering value.
This paper presents a design method for optimal guidance laws (OGLs) with multi-constraints when time-varying parameters are present in kinematic equations. This method uses cross-iteration to obtain time-varying parameter values and the piecewise function technology to solve the two-point boundary value problem. The proposed OGL can optimize the combination of landing angle, miss distance and control energy consumption. Ballistic simulations are conducted, and the landing angle using the OGL is twice more than using the proportional navigation law. The strike accuracy and damage effects increase because of the steep terminal trajectory. The time-varying parameters have no limit and require no simplification. Therefore, the proposed design method has a wide range of applications.
Ducted fans have higher thrust performance, higher propulsion efficiency and lower noise characteristics due to their duct system compared with commercial isolated propellers. The purpose of this study is to present design procedures in order to improve the aerodynamic performance of the ducted fan for the small VTOL UAV propulsion. In addition, the duct effect of ducted fans is analyzed to satisfy design requirements and improve performance. Aerodynamic design of the rotor and the stator blades of the ducted fan involves a series of steps: meanline analysis, through-flow analysis and aerodynamic analysis, based on consideration of the design requirements. The thrust performance of the ducted fan is somewhat higher compared with that of the rotor only, but wind tunnel test results of the ducted fan do not satisfy the design requirements. The thrust performance of the ducted fan is significantly different in the CFD results and wind tunnel results due to the inconsistency of the intake and the duct shape. Therefore, the thrust performance of the ducted fan would be somewhat improved by the optimization of the intake and the duct shape.
To investigate the ignition and combustion performance of the powdered fuel ramjet, an experiment is conducted on a connected-pipe ramjet test system. A computer code is developed to assist the design of the experimental powdered fuel ramjet. Based on the results of the parametric studies, an experimental powdered fuel ramjet is designed. The magnesium powder is pneumatically injected with ram-air. The flux of a two-phase mixture of solid powder suspended in the carrier air is measured with a Coriolis mass flow meter. Both cold-flow evaluation of the fuel feed system and hot-fire engine tests are performed. Ignition is realized successfully with a high-temperature combustion gas and high-energy spark plug. The feasibility of self-sustaining combustion of powdered fuel is demonstrated by introducing different flame holding techniques into the combustor. The multiple start-ups of the engine are also successfully performed using the spark plug as the ignition source. Methods for optimizing the combustion characteristics of the powdered fuel ramjet are put forward.
Blade vortex interaction (BVI) noise is a main source of helicopter noise especially during descending flight. Precise prediction of flow field around helicopters is required to establish the design technology of BVI noise reduction devices. Full CFD analysis including the rotor wake domain can predict flow field precisely, but it is computationally expensive. On the other hand, the prescribed wake model, which empirically predicts rotor trailing vortices, reduces computational cost greatly, but has less accuracy especially around the blade. Therefore, a hybrid method of CFD for the domain around the blade and the prescribed wake model is considered as a practical computational method in terms of the trade-off of computational accuracy and time. The base CFD code herein assumed is a structured grid Euler solver, 〈rFlow3D〉, which has been intensively developed for helicopter applications at the Japan Aerospace Exploration Agency. In this study, computational accuracy of the hybrid method is improved by applying circulation distribution, induced velocity distribution, and blade tip vortex position obtained in CFD domain to the prescribed wake model. The normal force coefficient on blade and noise contour computed by the hybrid method show good agreement with the experiment.
The present study attempts to predict the performance of hybrid rocket motors using a physics-based comprehensive model in which all the submodels that govern the various processes that occur during heating and combustion of solid fuel are based on physical-based mathematical models. Based on the fluid-solid coupling technique and some comprehensive physical processes during operation of hybrid rocket motors, a numerical model is developed to predict the regression rate for the solid fuel surface of the hybrid rocket motor under different operation conditions. The muti-dimensional Favre-averaged compressible turbulent Navier-Stokes equations are used as the governing equations of the reacting flow, the two-equation turbulence model is used to simulate the turbulent flow, and the eddy breakup model is used to simulate the gas combustion. The results presented are for hydroxyl-terminated-polybutadiene fuel and gaseous oxygen. The Navier-Stokes model results allow more detailed and realistic observation of the chamber flow field than is permitted by simpler boundary layer analyses. The model predications indicate that fuel surface regression rates are considerably impacted by both the size and geometry of the configuration. This study provides considerable information to the understanding of flow and combustion process in hybrid rocket motors. Methods for optimizing the hybrid combustion characteristic are put forward.
This paper deals with microsatellite attitude determination systems which are a combination of a main estimator with high-power, high-precision sensors for higher accuracy estimation and a redundant estimator with low-power, lower-precision sensors for backup. Measurement data from all sensors in the redundant estimator are fused by the unscented Kalman filter to provide estimated attitude and the gyro bias values. Besides the accuracy of attitude sensors, the accuracy of this estimator depends largely on the selection of the process and measurement noise covariance matrices. In this paper, a novel real-time tuning unscented Kalman filter for redundant attitude estimator is introduced to tune these matrices efficiently in each filter step. The tuning process uses the estimated attitude of the main estimator as an independent truth reference data to calculate the cost function which is minimized by a downhill simplex algorithm. In the scheme developed in this paper a fine-tuning process is used, which results in faster convergence speed and higher estimated accuracy of the redundant estimator. Another important feature of the developed filter is that a flexibly estimated accuracy and system power consumption can be archived by choosing the duration and repeat frequency of turn-on time of the main estimator.