Journal of Aerospace Engineering

The plane extension method was used to model aggregate shapes of three-dimensional (3D) concrete. In this method, the convexity condition requires that the volume formed by all the extension points and border planes be positive. A test method is adopted that involved summing the aggregate volumes to judge the intersection and overlap of aggregates. In the 3D random aggregate simulation (3D-RAS) structure, the shape, size, and spatial distribution of the aggregate particles resemble real concrete in the statistical sense, and fundamental aggregate diameters within the same gradation are assumed to follow random Gaussian distributions. The fracture process test of a three-point notched beam considering the effects of inhomogeneous material was used to validate the proposed aggregate model. Numerical illustrations show that aggregate shapes play an important role in the failure process of beams. Parameter studies also show that, although the homogeneity index has only a small influence on the elastic modulus, it plays an important role in the nonlinear softening behavior. The aggregate distribution has a little influence on the overall mechanical response of the concrete composite. The interfacial transition zone (ITZ) thickness has a great effect on the ultimate load of the concrete. An optimal aggregate volume fraction should exist for the concrete composites.

Geotechnical properties of the lunar regolith are critical parameters in the design of equipment for in situ resource utilization (ISRU) on the Moon. It is imperative to simulate the geotechnical behavior of the lunar soil properly in the development of such equipment. Soil strength depends on the preparation method as well as the properties of the granular raw material. To execute the many small- and large-scale equipment tests planned for ISRU, it is necessary to develop a simulant that is inexpensive and can be produced in large quantities. This paper presents the methodology behind developing such a lunar-like geotechnical soil, GRC-3, and compares the properties of this soil with those of lunar regolith. The results show that particle size distribution, specific gravity, bulk density, and shear strength parameters for several preparation protocols are similar to that of lunar soil. Therefore, GRC-3 can be used in future large-scale experiments to predict the performance of ISRU equipment on the Moon.

In the present paper, the solution for peristaltic flow of a six-constant Jeffreys fluid model in an endoscope is discussed. Perturbation and numerical solutions are calculated for nonlinear differential equations. A comparative study is also performed to examine the validity of both results. The results for pressure rise frictional forces and pressure gradient are calculated using numerical integration and presented graphically to interpret the behavior of various physical quantities of the six-constant Jeffreys fluid model.

Nonlinear longitudinal aerodynamics of planar fins and aircraft at high angles of attack near stall have been modeled using Kirchhoff’s formulation. Both the steady-stall and quasi-steady stall models have been applied on the longitudinal aerodynamic data at high angles of attack pertaining to planar fins and Hansa-3 aircraft, respectively. The National Wind Tunnel Facility (NWTF) in the department of Aerospace Engineering at the Indian Institute of Technology, Kanpur was used to generate the wind tunnel data for planar fins with rectangular and NACA-0012 airfoil cross-sections whereas flight test data near stall for Hansa-3 aircraft were generated by executing a quasi-steady stall maneuver (QSSM). The Kirchhoff’s steady-state stall model was applied to wind tunnel data for modeling the flow-separation point and to estimate the parameters characterizing stall. The Kirchhoff’s quasi-steady stall model was applied to the compatible real flight test data to model the nonlinear longitudinal aerodynamics and estimate the aerodynamic parameters. An approach of fixing the parameters to their reasonably accurate values was followed during the estimation of quasi-steady stall characteristics and the longitudinal aerodynamic parameters using the maximum likelihood (ML) method from the nonlinear flight data pertaining to QSSM. The results obtained using both models were found to be encouraging and are presented in graphical as well as in tabular form. The estimated parameters using both models have been compared with the reference results. A well-matched comparison of the model’s estimated and real flight data hysteresis loop is also presented.

Rotary wing aeroelasticity is a highly complex phenomenon involving coupling between flexible blade dynamics and unsteady aerodynamics including stall and unsteady wake effects. In this paper, a low-cost computational aeroelastic model including the structural coupling from geometric parameters and nonlinearities associated with structural modeling and dynamic stall, applicable to steady, level forward flight, has been developed. The differential equations of motion are solved in time domain in a sequential manner to obtain the response of all the blades in the rotor system, the dynamic inflow variables, and the sectional loads at every time step. A fourth-order Runge-Kutta integration scheme has been adopted for solving the differential equations. Iterations are carried out until convergence is achieved in blade response and helicopter trim. The effect of blade geometric parameters such as pretwist, hinge offset, and torque offset on aeroelastic response of a helicopter rotor system is investigated numerically. It is shown that the structural coupling from blade geometric parameters significantly influences the rotor blade response and loads.

Motion planning determines trajectories for vehicles that link an initial location and heading with a final location and heading. Techniques for motion planning have been developed for two-dimensional maneuvering; however, they are less mature for three-dimensional maneuvering. The concept of motion primitives is particularly attractive for motion planning that determines trajectories as a set of maneuvers that satisfy differential constraints. This paper furthers work on a higher-level abstraction of trajectory primitives that consider sequences of motion primitives. In this paper, trajectory primitives are developed that deal with airspace constraints of an environment. The motion planning is shown to be an optimization involving a pair of trajectory primitives that is related by an intermediate waypoint. The resulting path is completely parameterized by the waypoint location.

In this paper, the aeroelastic modeling and flutter analysis of swept aircraft wings carrying two powered engines is considered. The aeroelastic governing equations are determined via Hamilton’s variational principle; the equations include both concentrated mass and the thrust force terms. Also, Peter’s finite-state aerodynamic model, modified to take the sweep effects into account, is employed. The effect of engine thrust is modeled as follower forces of prescribed magnitude. Furthermore, flutter speed predictions are validated with the published results and good agreement is observed. Four different statuses of the engine thrust are selected to cover all possible situations in realistic cases. In the first case, both engines are turned off and the wing is simply affected by two external stores. The second and third cases correspond to the situation in which aircraft wing is modeled with one powered engine. This can take place when one of the engines is out of order or switched off. Finally, in the last case, two powered engines are investigated. The effects of the wing sweep angle and other design parameters such as the mass ratio and engine locations on the flutter boundaries are presented for these situations. The results show that the thrust and mass of the two engines has a complicated influence on the flutter boundaries.

The picosatellite for remote sensing and innovative space missions (PRISM) is a 8.5-kg nano remote sensing satellite that had been launched in 2009. The objective of the mission is to obtain 30-m resolution Earth images using an extended optical system. To achieve the requirement, the satellite attitude should be stabilized to an accuracy of $0.7\text{°}\text{/}\text{s}$. The focus of this research is the attitude determination and control system for the PRISM satellite. The satellite controls the angular velocity using gyro sensors, magnetometers, and magnetorquers. For an accurate attitude control system, the satellite should calibrate gyro sensors and magnetometers. Generally, attitude sensors are calibrated using ground calibration systems or more precise attitude sensors in-orbit, whereas a nanosatellite is difficult to use with these calibration systems and precise sensors because of high cost. The PRISM satellite estimates the scale factor and bias of the magnetometers with the International Geomagnetic Reference Field model. Then, the satellite estimates the bias and scale factor of the gyro sensors with calibrated magnetometer measurements. After the calibrations, the satellite stabilizes the attitude with the cross product control method. The paper presents a method to calibrate attitude sensors and to stabilize a nano remote sensing satellite attitude.

Wind disturbance is one of the major factors that could jeopardize a safe landing. To improve the safety of automatic landing, in this paper, a linear model of the aircraft in longitudinal motion is established first and further augmented by introducing the tracking error integral equations to achieve a desirable tracking performance. Then, two different design methods are proposed to obtain robust controllers in the presence of parametric uncertainties, external disturbances, and input constraints. One of the methods uses the $H_{\mathrm{\infty }}$ state-feedback and regional pole placement technique, and the corresponding controller is designed via a Lyapunov approach. The other method develops the robust exact pole placement technique subject to the input constraints, under which framework the controller is designed through minimizing the condition number of the solution of a Sylvester-like equation. The existence conditions for both admissible controllers are formulated as linear matrix inequalities. Finally, the Boeing 747 airplane and the windshear model are employed to show the effectiveness and robustness of the proposed design techniques.

Nanospacecraft are usually launched in clusters from a single launch, in which they are hosted as secondary payloads. The nanospacecraft, e.g., CubeSats, are usually released simultaneously in small groups from one single container, to simplify the launcher interface manufacturing and the launch operations. The release sequence must be designed to minimize the risk of collisions among the nanospacecraft themselves. Considering the increasing number of micro- and nanosatellites in recent space activity, a collision risk analysis and a release systems configuration optimization are of interest. This paper describes a collision probability analysis for a nanosatellite cluster deployment, performed by a series of Monte Carlo simulations, comparing results obtained with different release mechanisms, procedures, and orbital dynamics modeling assumptions. The collision risk in the first orbits after launch depends mostly on the in-orbit injection initial conditions, such as relative initial position and velocity, which directly reflect on the satellite dispenser configuration, interface concept, and mechanical parameters. The model used for the analysis is based on the Encke equations for the relative motion, considering the main perturbations acting on the satellites, such as Earth gravitational field higher harmonics, Moon and Sun third-body perturbations, solar radiation pressure, and atmospheric drag. The risk of collisions is assessed performing Monte Carlo simulations based on the numerical integration of the equations of motion. The final results provide the collision probability assessment and the influence that the release device configurations have on this risk.

The unscented Kalman filter (UKF) is a filtering algorithm that gives sufficiently good estimation results for estimation problems of nonlinear systems even when high nonlinearity is in question. However, in the case of system uncertainty the UKF becomes inaccurate and diverges in time. In other words, if any change occurs in the process noise covariance, which is known a priori, the filter fails. This study introduces a novel adaptive fading UKF algorithm based on the correction of process noise covariance ($\mathbf{Q}$-adaptation) for the case of mismatches with the model. By the use of a newly proposed adaptation scheme for the conventional UKF algorithm, change in the noise covariance is detected and corrected. Differently from most of the existing adaptive UKF algorithms, covariance is not updated at each step; it has only been corrected when the change in the process noise covariance is detected, and that brings about a noteworthy reduction in the computational burden. The proposed algorithm is tested as a part of the attitude estimation algorithm of a picosatellite, a satellite type for which computational convenience is necessary because of the design limitations. In this sense, different change scenarios for the process noise covariance are taken into consideration.

This work concerns a collision-free fixed-time formation reconfiguration control method for unmanned aerial vehicles (UAVs). The reconfiguration to the new formation in three-dimensional space is specified based on the target final states of each UAV. A reference reconfiguration trajectory is generated from a Bolza optimization solution of simplified dynamics of the vehicle. A sliding controller is then utilized to track the reconfiguration trajectory. Collision avoidance is achieved by modeling the detection region of each UAV as a potential field. Such a field generates control signals that are inversely proportional to the square of the distance between the UAV and a moving or stationary object within its detection region. Altered optimal trajectories are generated online as the UAV avoids collision with another UAV or other obstacles. Simulations confirm that the control scheme developed is able to produce satisfactory results. This paper examines the stability (in the sense of Lyapunov) and performance of the controller under the influence of the avoidance control command. The results of this study provide useful insights to the behavior of the UAV formation under tracking command and a collision avoidance system.