Journal of Aerospace Engineering

In this paper, the authors tried to design the altitude hold mode autopilot for unmanned aerial vehicles. This case presents an interesting challenge attributable to the nonminimum phase characteristic, nonlinearities, and uncertainties of the altitude to elevator relationship. A fuzzy logic autopilot in a single-loop scheme is proposed for the design of this autopilot. The multiobjective genetic algorithm is used to mechanize the optimal determination of fuzzy logic autopilot parameters on the basis of an efficient cost function that comprises undershoot, overshoot, rise time, settling time, steady state error, and stability. Simulation results show that the proposed strategy not only has a simple structure, but also has desirable performances in time response characteristics, robustness (against the unmodeled nonlinear terms and parametric uncertainties), and the adaptation of itself than the large commands.

The formation and development of a wing-tip vortex in a near and extended near field were studied experimentally. A swept-back tapered wing with a NACA23012 cross-section was used as a vortex generator. Particle image velocimetry (PIV) as a whole field velocity measurement technique was used in a low-speed wind tunnel to measure and characterize the wing-tip vortex. Wake structures at successive downstream planes crosswise to the axis of the wake vortices were evaluated in internal and external core radius, maximum tangential velocities, vorticity, and circulation distributions. The effect of angle of attack on vortex parameters was examined at one downstream location. Internal core radius and circulation distributions were nearly constant along the downstream direction. A direct dependence of the circulation and tangential velocity distribution on the angle of attack was evident. The centers of the wing-tip vortices scatter in a circle of radius nearly equal to 1% of the mean wing chord. Meandering amplitudes showed no direct dependence on the vortex strength but increase along the downstream direction. Good agreement was obtained between the theoretical exponential vortex model and the measured data. Computed induced rolling moment coefficients generated by the wing are within the full roll control capability of a follower aircraft.

Under the perturbations of the $J_2$ zonal harmonics term and air drag, to obtain a set of dynamics equations that have a simple and concise form and can be used to exactly propagate the relative motion of a chaser spacecraft with respect to a reference spacecraft in an arbitrary elliptical reference orbit, Newtonian Hamiltonian mechanics are adopted in the derivation process. On the basis of the principle of Newtonian mechanics, six first-order differential equations are derived to accurately describe the inertial motion of a reference spacecraft under the affect of $J_2$ term and air drag. By using the principle of Hamiltonian mechanics, another six first-order differential equations are obtained to precisely propagate the relative motion together with inertial motion dynamics equations of the reference spacecraft, without the equation of the right ascension of ascending node $\Omega$ . Because the spherical gravitational potential, the $J_2$ zonal harmonics gravitational potential, and air drag are axisymmetric about the Earth rotation axis and independent of the motion of $\Omega$ , the relative motion is independent of $\Omega$ . So $\Omega$ is unnecessary in the relative motion propagation. The affect of $J_2$ zonal harmonics term and air drag on precession of the reference orbital frame also is analyzed. On the basis of derivation and analysis, the dynamics equation of relative motion around an oblate Earth with air drag are promoted to four application areas: A1, long-time relative motion prediction; A2, field of view analysis for relative measurement sensor; A3, the analysis of affect of coordinate frame precession on reference spacecraft attitude control; A4, the analysis of affect of of coordinate frame precession on attitude synchronization mission. Simulation results in application (1) confirm the validity and exactness of the dynamics equations derived in this paper. All applications indicate that the dynamics equations have a good prospect of application in formation flying mission.

In the space attack defensive system, a satellite is usually used to attack another satellite. To approach the target satellite successfully, path planning is very important for the attacking satellite, especially when the defensive system of the target satellite is composed of formation flying small satellites. This paper presents an environment model that is established on the relative orbit model of small satellites and changed into a static model for path planning. A genetic algorithm is used to find the optimal path. The coding rules, fitness function, and three genetic operators are designed in turn. The fitness function is found by considering region constraint, safe constraint, shortest length constraint, and crooked degree constraint. The simulation results show that the optimal path can be found with good stability and convergence, and the optimal path could satisfy requirements of security and the shortest path well.

On-board black body is used for radiometric calibration of spaceborne infrared (IR) radiometers and imaging systems. The black body is equipped with heaters to heat up the black body during imaging for calibration, a heat pipe to transfer residual heat on the black body just after the black body heating to radiator on the spacecraft (S/C), and embedded temperature sensors to measure the black body surface temperature has been proposed and manufactured. To investigate the performance capability of the black body, functional tests were performed, such as emissivity measurement tests and black body heat-up tests under ambient conditions. The characteristic of black body under on-orbit conditions has been predicted by numerical simulation using a correlated thermal model obtained from ambient test results.

Flapping-wing vehicles produce aerodynamic lift and thrust through the flapping motion of their wings. The performance of a flexible-membrane flapping wing is experimentally investigated here. A new scheme for finding flight envelopes for a particular flapping vehicle is introduced as well. A flapping-wing system and an experimental setup are designed to measure the lift, thrust, and power usage of the flapping-wing motion for different flapping frequencies, angles of incidence, and various wind tunnel velocities up to $12\mathrm{m}/\mathrm{s}$ . The obtained results are used in the performance analysis of this flapping wing. For each given weight (or payload), one may find the best cruise speed for maximum range, corresponding to a minimum ratio of power to cruise speed (i.e., energy usage for unit distance), or for maximum endurance corresponding to minimum power (i.e., minimum flapping frequency).

In this paper, nonlinear flight dynamics of a flexible aircraft under aeroelastic and gust loads is presented. Hybrid state governing equations that incorporate both aircraft rigid-body motions and elastic deformations are utilized. The aircraft wings perform as classical beams, and the structural model, which incorporates bending–torsion flexibility, is used. In addition, quasi-steady and Peter’s unsteady aerodynamic pressure loadings are considered and modified to take the gust velocity components into account. The distributed parameters representing elastic deformations of flexible components are discretized by using Galerkin’s method. Three different excitations are considered for nonlinear simulation: aileron, elevator, and vertical/pitch components of Dryden gust model. Time responses of the rigid vehicle and elastic aircraft with quasi-steady and unsteady aeroelastic model are presented, and differences are highlighted.

In the present article, the effects of temperature-dependent viscosity on the peristaltic flow of Jeffrey fluid through the gap between two coaxial horizontal tubes was analyzed. The inner tube is maintained at a temperature $T_0$ , and the outer tube has sinusoidal wave travelling down its wall and is exposed to temperature $T_1$ . The governing problem is simplified using long-wave length and low Reynolds number approximations. Regular perturbation in terms of small viscosity parameter is used to get the expressions for the temperature and velocity for Vogel’s models of viscosity. The numerical solution of the problem has also been computed by a shooting method, and an agreement of numerical solutions and analytical solutions had been presented. The expressions for pressure rise and friction force are calculated numerically. Graphical results and trapping phenomenon is presented at the end of the paper to see the physical behavior of different parameters.

This paper presents attitude trajectory control for a satellite system. The pitch loop of the satellite is controlled by momentum wheel; whereas the roll/yaw loops are stabilized by two different magnetic torques along their axes. The controller is on the basis of both the linear quadratic regulator (LQR) control theory and recursive least square (RLS) algorithm that is used to identify the equivalent linear canonical second order system for each axis at discrete points on references. Then, the linear identified model is used to design the LQR-based controller. This process is done in each time span. The efficiency of the controller to bring the position error close to zero is shown by some numerical simulations. The performance of the proposed controller is remarkable in stabilizing the satellite attitude even in the presence of orbital eccentricity.

This study introduces a robust Kalman filter (RKF) with a filter-gain correction for cases of measurement malfunctions. Using defined variables called measurement-noise scale factors, the faulty measurements are taken into consideration with a small weight and the estimations are corrected without affecting the characteristics of the accurate ones. In this study, RKF algorithms with single and multiple scale factors are proposed and applied for the state estimation process of an unmanned aerial vehicle (UAV) platform. The results of these algorithms are compared for different types of measurement faults, and recommendations for their utilization are given.

The thermomechanical behavior of damaged space shuttle tile thermal protection system (TPS) is simulated using the finite-element method. The effects of damage on the thermal protection capability and the induced thermal stresses in the TPS are evaluated by comparing the thermal and structural response of the damaged configurations with the undamaged configurations. The TPS consists of three components (the LI-900 tile, the strain isolation pad, and the underlying structure), and it is modeled as a discrete, three-layer structure. The TPS is subjected to the reentry heating and pressure profile corresponding to the Access to Space vehicle. The transient temperature distribution and the resultant thermal stresses are then computed. Three different sizes of damage simulating hypervelocity impact are considered. The validity of the simplifying assumptions used in an earlier study is systematically examined to determine the correct combination of assumptions required for modeling this class of problems. Influence of damage location on TPS response is also studied and it is concluded that centrally located damage represents a worst-case scenario.

The use of epoxy adhesives in metal structures for adhesive reasons is a continuously growing market, because of the increasing demand of lightweight constructions, which renders in certain cases conventional mechanical bonding unpractical. Furthermore, epoxy adhesives possess enhanced mechanical, chemical, and physical properties, i.e., increased shear and compression strength and resistance in solvents, whereas maintaining these properties at higher temperatures than traditional mechanical bonding methods. Next generation aircraft engines and pipelines are common application areas of these adhesives. In order, however, to ensure the safe use of epoxy adhesives in such structures, computational analyses must be conducted to simulate eventual failure mechanisms. These simulations require the exact determination of the epoxy adhesion strength properties with respect to several parameters. A systematic investigation of these properties will be presented, on the basis of tensile tests of standard aluminium tension specimens bonded together with epoxy adhesives, whereas their failure mechanisms will be simulated by a FEM-supported model to evaluate the extracted characteristics. The specimens were cemented by means of a developed experimental device under constant temperature and humidity conditions. The deformation as well as the developed stress distribution of the specimens during the tensile experiments were simulated by means of a FEM-supported software package, considering the stress-strain curves of the adhesive, determined by a FEM-based evaluation of microindentation results.

Some in-service deterioration in any mechanical device, such as an aeroengine, is inevitable. As a result of experiencing deterioration (of the engine as a whole or any of its components individually), an engine will seek a different steady operating point, thereby resulting in a variation of the specific fuel consumption and/or fuel flow to provide the same thrust to keep the aircraft’s performance invariant. Any rise in specific fuel consumption and/or fuel flow and thereby the increased quantity of fuel required is of prime importance in military aviation. For a military aircraft’s mission profile, using computer simulations, the implications of high-pressure turbine’s erosion of a turbofan aeroengine upon the weight of the fuel that has to be carried and consumed are predicted. This will help in making wiser management decisions, such as whether to remove an aeroengine from the aircraft for maintenance or to continue using it with some changes in the aircraft’s mission profile). Hence, improved engine utilization can be achieved, resulting in lesser overall life-cycle costs.

Three axis-control low-earth-orbit (LEO) satellites use momentum exchange devices, such as reaction wheels, to maneuver and control their attitude to perform image acquisition. High-resolution satellite payloads demand stable pointing with very small jitter to achieve good performance. Microsatellites with flexible appendages such as deployable solar panels are very prone to vibration, especially when excited by internal actuators such as reaction wheels or momentum wheels that are used to control the attitude of the satellite. Any infinitesimal amount of unbalance in the reaction wheels’ rotors will impose a harmonic excitation that may interact with the structure of the solar panels. Therefore, controlling the solar panels’ structural vibration interaction would control the jitter. This control was achieved by attaching a set of properly chosen point masses at selected locations along the solar panels. The panels are modeled using Timoshenko beam theory. A scheme was proposed for choosing the optimal values of the added masses and their locations to minimize the satellite vibration. A prototype of a microsatellite with three reaction wheels and two solar panels, equipped with linear actuators to control the point mass location along the panels, is developed to demonstrate the concept.

Twenty-one swing tests were conducted at the NASA Langley Landing and Impact Research (LandIR) Facility in 2006 to evaluate Orion boilerplate ground landings as a function of impact velocity and pitch. In this article, experimental results from two capsule swing tests and three vertical drop tests are compared with numerical results. The numerical results were derived by using the nonlinear dynamic finite-element code LS-DYNA. The swing tests were conducted at 0° and $-15°$ pitch, impacting a level soil mat at a $17.68\mathrm{m}/\mathrm{s}$ (58 fps) horizontal velocity and $1.52\mathrm{m}/\mathrm{s}$ (5 fps) vertical impact velocity. Three vertical drop tests were conducted with the boilerplate capsule at 0° pitch. Each test was conducted to measure the impact on a level soil mat with varying impact velocities: $3.66\mathrm{m}/\mathrm{s}$ , $7.31\mathrm{m}/\mathrm{s}$ , and $10.97\mathrm{m}/\mathrm{s}$ (12, 24, and 36 fps). Results of the study show the potential of using numerical modeling for vertical drop test and swing test simulations. Discrepancies between the numerical and experimental simulations primarily stem from the nonuniformity and complexity of soil behavior.

To realize a long-term lunar exploration, it is essential to develop a technology for transporting lunar soil for in situ resource utilization. We are developing a particle transport system that utilizes electrostatic traveling waves. The conveyer consists of parallel electrodes printed on a plastic substrate. Four-phase rectangular voltage is applied to the electrodes to transport particles on the conveyer. Mechanical vibration was applied to the conveyer to transport particles more efficiently. The results of our investigation are as follows: (1) The observed transport rate in air was $13.5\mathrm{g}/\min$ for a conveyer with a width of 100 mm. By performing numerical calculations on the basis of the 3D distinct element method, we predicted that the system performance would improve in the high-vacuum and low-gravity environment on the moon. (2) Power consumption in this system is much less. It was only 10 W for a conveyer with an area of $1.0{\mathrm{m}}^2$ . (3) We demonstrated an inclined and curved transport path as well as a flat and straight transport path. In addition, we demonstrated that transportation of particles through a tube and accumulation of scattered particles were also possible.

The removal of lunar dust adhering to spacesuits is essential for the long-term exploration of the Moon. Because lunar dust is magnetic, a convenient cleaning device that utilizes magnetic force has been developed. The device can capture and separate lunar dust continuously with a multipole magnetic roller. The observed separation rate was 90%, but the capture rate was low, resulting in an overall cleaning rate of 40%. The rate increased when large amounts of dust adhered to the fabric. Although it was difficult to capture small particles trapped between fabric fibers, this device will be able to preliminarily clean spacesuits and save the precious activity time of astronauts on the Moon.

A further improvement is attempted of an existing analytical model for an accurate prediction of the aeroelastic stability of a tiltrotor aircraft. A rigid-bladed rotor structural model with the natural frequencies selected appropriately in both the flapping and lagging motions is used. The geometric coupling between the wing vertical bending and torsion is also included. The pitch-flap and pitch-lag couplings are also added. Three different aerodynamic models are combined with the structural model: two quasi-steady models and one full unsteady aerodynamics model. Frequency domain analysis is conducted to predict the whirl flutter stability boundary. It was found that the geometric coupling must be included at an appropriate level to predict the whirl flutter boundary accurately. The addition of the wing bending/torsion coupling and the control system flexibility improves the prediction accuracy significantly. Unsteady aerodynamics influences the stability prediction. The whirl flutter boundary is predicted to be less when quasi-steady aerodynamic models are used as compared with that for unsteady aerodynamics. For these two cases, different behaviors regarding the intersection among the relevant structural modes, especially between the lower frequency rotor modes and the wing modes, are observed from the frequency and damping prediction.