For the development of an unmanned aerial vehicle (UAV) simulator, an aerodynamic analysis was performed. To analyze the aerodynamic coefficient, an unsteady numerical method (time-domain panel method) was used. This panel method was based on the Dirichlet boundary condition coupled with a time-stepping method. The Free Wake method was also used to get more accurate results. The proposed numerical method was validated by comparing our results with experimental data. Through the aerodynamic analysis, longitudinal, lateral, directional and dynamic derivatives were obtained. To build the database for the UAV, the results were predicted rationally. This research will contribute to designing UAVs as well as to developing UAV simulators.

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The cantilevered ramp injector has been employed as an effective fuel injection strategy in the shock-induced combustion ramjet engine, and its flow field characteristics have attracted increasing attention. Three-dimensional Reynolds-Averaged Navier-Stokes (RANS) equations coupled with the RNG k-ε turbulence model have been employed to investigate the flow field of a cantilevered ramp injector in a freestream with a Mach number of 6.0. The effects of the angles of attack and sideslip on the flow field of a cantilevered ramp injector have been studied, and the predicted axial distribution of the maximum injectant mole fraction shows good agreement with the available experimental data in public literature. The obtained results show that the case with an angle of attack 3° can promote the mixing process between the fuel and air most efficiently in a typical cantilevered ramp injector. However, it cannot prevent the premature ignition of the premixed combustible flow. The mixing process cannot be promoted when the normalized axial distance is larger than 7.55 in cases with different angles of sideslip, and larger angles of sideslip can reduce the effect of premature ignition for the premixed combustible flow.

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This paper presents spacecraft orbit estimation using only the geomagnetic field measurement via a batch processor based on the unscented transformation. Orbit estimation is performed for different inclinations and observation spans. Truth model based tests show that the magnetometer-based orbit determination has a dependency on the orbital inclination; i.e., on the gradient of the geomagnetic field. This is the main reason of the dominant cross-track error in a polar orbit and the along-track for an equatorial orbit. This phenomenon is mainly caused by the dipole property of the geomagnetic field. Additionally, the proposed unscented batch filter is compared with the Bayesian batch filter to investigate the filter's performance. The comparison results show that the unscented batch filter has strength in convergence speed over the Bayesian filter. The proposed algorithm is evaluated using the real magnetometer data of the CHAMP. The achieved position accuracies are similar in the both filters; i.e., approximately 1–2 km depending on the magnetometer noise level. However, the unscented batch filter can estimate the orbit more rapidly than the Bayesian filter when nonlinearity is strengthened. It shows the unscented batch filter has strength beyond that of the Bayesian one in highly nonlinear situations.

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The integrated missile design optimization process is proposed by implementing the aerodynamics database (Aero DB) and tactical missile design (TMD) spreadsheet to obtain a quick and relatively accurate optimal air intercept missile configuration at the conceptual design stage. The Aero DB is constructed to replace an existing aerodynamics analysis module in the TMD spreadsheet and to provide stability and control coefficients as constraints for improving missile range performance based on the body-wing-tail configuration baseline. Sensitivity analysis is performed on an entire missile geometry and flight condition variables to eliminate the small effects of design variables on missile range and constraints under a PHX ModerCenter^® 10.1 integration environment. The optimal missile configuration shows 27.8% improvement in total range compared with a body-wing-tail configuration baseline while all constraints are satisfied. The proposed integration of the missile design program using Aero DB demonstrates more accurate and reliable results which are validated by high-fidelity analysis ANSYS Fluent 13^® on the optimal missile configuration compared with TMD aerodynamics analysis results. The maximum difference between ANSYS Fluent and Missile DATCOM is 11.76% at 10 degrees of AoA compared with 37.97% for TMD aerodynamics analysis and ANSYS Fluent difference.

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The large thermally induced bias drift is the main factor affecting the performance of the interferometer fiber optic gyroscope (IFOG) in practical engineering applications. The thermally induced bias drift is investigated in order to improve the accuracy of IFOG in the strapdown inertial navigation system (SINS). A triaxial integrated linear multi-variable model (TILMM) is presented based on experiments. The model is composed of the temperature gradient, polynomial temperature and temperature square. Due to the linear character of the model, model parameters are identified from the temperature and system outputs by least-square (LS) estimation methods. The drift data sets are collected to validate the effectiveness of this model. The rough experiment data is preprocessed by the proposed sliding window median absolute deviation (SWMAD) technology to identify outliers firstly, then the gradient inverse weighted (GIW) filter is used to initially lower the noise, and the singular value decomposition (SVD) is used to reduce the noise of the experiment data. The verification experiment of compensation is conducted and shows that the TILMM is effective, and SINS bias stability is improved in comparison with single-axis compensation.

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A new control scheme of nonlinear output tracking and disturbance rejection for autonomous close-range rendezvous and docking of spacecraft is proposed in a closed-loop structure. The robustness of the proposed control scheme is achieved through the sum of output-feedback control and state-feedback control as a closed-loop control structure. The designs of both control laws are based on full, 6-degrees-of-freedom, nonlinear dynamic models, which are transformed into a linear-like form using a state dependent Ricatti equation technique. The proposed control scheme ensures close-range rendezvous and meets the state tracking conditions for docking such that the convergence of state tracking errors should be ensured in the presence of unknown disturbances and uncertainties of the system parameters. Compared with the conventional state-dependent Ricatti equation method, the proposed control scheme yields superior state tracking performance and robustness in the presence of external disturbances and uncertain system parameters. The 6-degrees-of-freedom numerical simulation results are presented to verify the effectiveness and capability of the proposed control scheme for autonomous close-range rendezvous and docking.

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To use the L₂ point of the Earth-Moon (EM) system as a space transportation hub, it is absolutely necessary to know the available orbital energy of a spacecraft departing from EM L₂ and how to obtain required excess velocity to the target planet. This paper presents a successful acceleration strategy using the combination of an impulsive maneuver and the Sun gravitational perturbation. Applying this strategy enables the spacecraft to efficiently convert the energy of the maneuver into its orbital energy.

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