Proceedings of the Institution of Mechanical Engineers,Part G: Journal of Aerospace Engineering

In this article a L 1 adaptive state feedback controller is presented for the three-dimensional integrated guidance and control of interceptor. The model of three-dimensional integrated guidance and control is first built by using engagement kinematics and interceptor dynamics. The objective is to control the fin deflection angles to make the line-of-sight rates converge to zero. The L 1 adaptive control is then applied to design the three-dimensional integrated guidance and control controller. The L 1 adaptive controller guarantees uniformly asymptotic and bound transient tracking for the system inputs and outputs, which are important for interceptions with strong robustness and high timeliness. The design issues of L 1 adaptive controller are also discussed. In the interception simulation, the L 1 adaptive state feedback controller demonstrates the robustness to different kinds of uncertainties, while guaranteeing the transient performance of dynamic interception process.

This article aims to compensate the velocity and position errors that exist when the star sensor starts to work in a strapdown inertial navigation system aided by celestial navigation. These systems are integrated via unscented Kalman filter to estimate the current attitude and the gyros fixed bias, precisely. Since an accurate integration is desired, the nonlinear attitude equations are utilized in filter and these equations are propagated through a precise discretization method. Then, implementing the back-propagation and smoothing techniques, the initial attitude and the accelerometers fixed bias are also estimated. Finally, carrying out a parallel navigation, the velocity and position errors are compensated. The validity of the proposed method is investigated through simulation of launch vehicle navigation. Simulation results show a great reduction in velocity and position errors.

This article proposes a new strategy that computes the bank-to-turn commands without a singularity problem. To this end, the singularity problem is first analysed, and the main influence factors are found. An extended roll-angle command calculation method is then derived for the missile body coordinate based on the bank-to-turn-90 logic. The auxiliary skid-to-turn manoeuvring and the command increment saturation are induced to eliminate the oscillation of roll-angle command due to the noises in guidance commands. Three control zones are designed to ensure that suitable command calculations are for different conditions. When the strategy is used, the missile tends to maintain a smooth and varied roll-angle command, even if the guidance acceleration commands approach zero at the endgame of guidance. Finally, numerical simulation results are provided, and the validity of the strategy is proven via a comparison between the typical bank-to-turn guidance law and the normal bank-to-turn command calculation method.

For large impact angle control problem (here, the “large impact angle” means the impact angle in the closed interval from −180° to 180°), estimating the time-to-go accurately is the key of impact time and impact angle control guidance (ITIACG). The objectives of this paper are to construct a new impact angle control guidance (IACG) law suitable for large impact angle control and present a time-to-go estimation procedure for the new IACG law suitable for designing ITIACG law. The constructed IACG law is a biased proportional navigation guidance law with large impact angle constraint, the rule of the cosine of the lead angle in the biased term is to guarantee that the lead angle remains in the open interval from −90° to 90°, which is required in the development of time-to-go estimation procedure. To estimate the time-to-go, by introducing a self-convergent angle named as alfa, the closed equations of motion are transformed to a different form, which can be solved conveniently under the assumption of small lead angle. For the case of large lead angle, the time interval of time-to-go is partitioned into n segments, the maximum increment of lead angle is supposed to be a small angle in each segment, the transformed closed equations of motion can be expressed as function of alfa angle and solved analytically. A geometric approach is proposed to determine conservatively a suitable alfa angle to guarantee that the maximum increment of lead angle is a small angle in each segment. The time-to-go estimation procedure for the new IACG law are illustrated. Simulations are performed to verify the effectiveness of the proposed IACG law and the accuracy of the time-to-go estimation procedure.

Current Air Traffic Management research programs (i.e. Single European Sky ATM (air traffic management) Research (SESAR), next generation air transportation system (NextGen)), try to overcome airside capacity shortages while improving cost-efficient operations and safety. An increment in the airspace traffic density can lead to congested traffic scenarios for which it becomes necessary to develop new safety procedures that address multithread threats. This paper considers some of the difficulties in establishing validation of the airborne collision avoidance system (ACAS), which constitutes the last-resort for reducing the risk of near mid-air collision between aircraft in a multithread scenario. A causal model that is specified in Colored Petri Net (CPN) formalism is presented as a key approach to analyze the state space of a congested traffic scenario in which the events that could transform a conflict into a collision are identified, providing a challenging tool not only for validation but also for the implementation of a new ACAS logic. The InCAS EuroControl simulator has been used to illustrate the importance of cause–effect analysis for the relationships between various encounters that arise in a multithread scenario, in which TCAS II v. 7.1 fails to avoid a collision when two Resolution Advisories are issued without consideration of the downstream effects.

This paper presents a combinatorial optimization method based on uniform design in combination with response surface methodology and genetic algorithm. Uniform design is used to obtain experimental points and response surface methodology to establish a mathematical regression model. Subsequently, genetic algorithm is employed to acquire optimal solution of the objective function. The optimization method has been applied to a two-dimensional S-shaped transition duct design. The process is performed with two design variables. One defines the drop height ratio which describes wall profile, and the other depicts the length ratio between the axial length of the S-shaped transition duct and the duct inlet height. Total pressure loss coefficient as an aerodynamic performance parameter is selected as the objective function for optimization. The objective function is numerically assessed at design points sampled by uniform design in the experimental domain. The initial transition duct was designed with a radius-change to length ratio 11.6% larger than current engine design limits, and the optimization yields a decrease of 36.9% in total pressure loss and more uniform distributions of parameters at the outlet. The paper shows that the described optimization method can be applied to turbofan engines to increase the radial offset and decrease the axial design space between the fans and cores without jeopardizing performance.

To improve the performance of aero gas turbine engines, more and more interests have been shown on turbine inter-guide-vane burner based on the ultra-compact combustion concept. To make a universal turbine inter-guide-vane burner, a new concept is proposed using a trapped vortex cavity to replace the high swirling circumferential cavity combustor to address the need to scale the configuration for a larger turbine spool. Three models, including trapped vortex combustor, transition model, and turbine inter-guide-vane burner, are designed. Comparative analysis between combustion performances of three models by using numerical simulation method is carried out. The scale-adaptive simulation turbulence model is used in the simulation process, aiming to reduce the deviation between numerical simulation value and actual value. Finally, the turbine inter-guide-vane burner model is found to be the superior design proposal for turbine inter-guide-vane combustion technology, compared with the other two models.

The force equalization of a hybrid actuation system combining one servo-hydraulic actuator and one electro-mechanical actuator operated in position control and in active/active mode is addressed for safety critical applications such as primary flight controls. In a first step, an accurate virtual test bench is built to facilitate the analysis of force fighting and the assessment of the performance and robustness of the proposed force equalization strategies. It is validated from real experiments performed for the aileron actuator of a single-aisle commercial aircraft. Static force equalization is achieved first by adding equalization offsets in the position control loops as a function of the integral of the force difference between actuators. In order to keep a high level of segregation, the authority for this action is limited to 4% of the total actuator stroke. The dynamic force equalization is performed by forcing the two actuators to follow the same path. Thus, a trajectory generator is introduced to output the required position, velocity and acceleration from the position set point with realistic reproduction of the actuator power limits. Feedforward actions are used to compensate the major and invariant effects such as servo-hydraulic actuators functional flow and electro-mechanical actuator inertial torque. In this way, the pursuit errors are significantly reduced without decreasing robustness. Then, the accurate virtual test bench is used to assess the robustness of the force equalization strategy by analyzing the sensitivity of performance indicators to parameters and operating conditions. It is shown that the proposed force equalization scheme meets all the requirements, including segregation, robustness and simplicity.

In this study, a general assessment of inverse trigonometric shear deformation theory, recently developed by the authors, is performed and the structural responses (static, buckling, and free vibration) of laminated-composite and sandwich plates are investigated. The in-plane displacement components are expressed in terms of an inverse cotangent function, which yields the nonlinear shear deformation while the constant transverse displacement is assumed over the thickness of the plate. A computationally efficient finite element model for the implementation of above-mentioned theory is proposed. The continuity requirement of the finite element model is maintained as C⁰ by a suitable choice of nodal field variables. Numerous analysis problems are selected to study the effects of various parameters such as span-to-thickness ratio, lamination sequence, loading conditions, boundary conditions, etc. on the response characteristics of plates. Higher modes are also obtained for the buckling and vibration problems and the ability to investigate higher modes is ensured. The comparison of the present results with the established results in literature indicates the efficiency and range of applicability of the present formulation. Moreover, the formulation is presented in a generalized approach which enables the implementation of all existing seven degree-of-freedom theories in a single computer algorithm thereby making it practically more significant.

The aim of this paper is to account for the effect of the epistemic uncertainty of the input variables’ uncertainty in the nonprobabilistic reliability analysis on the safety of the structure system. Based on the idea of moment-independent sensitivity analysis, a modified sensitivity measure of the nonprobabilistic reliability is constructed to identify the most influential epistemic parameters of interval variables. For calculating the nonprobabilistic reliability sensitivity measures of the epistemic variables, a computational model is established. And a solution method with the advantages of the state-dependent parameter model is employed to improve the computational efficiency and avoid the complex sampling procedure. The numerical examples and engineering examples show that the proposed method of solving the sensitivity measure is reasonable and effective. The sensitivity measure of nonprobabilistic reliability proposed in this paper can give an essential importance sequence of all the epistemic uncertainties and identify key contributing epistemic uncertainties. When the sensitivity measure is larger, the epistemic uncertainty variable will become more important and should collect the data to increase knowledge of parameters. The sensitivity measures can provide the availability guidance to reduce the number of epistemic variables.

In geomagnetic aided navigation, directional matching suitability can be depicted by the directional features extracted from candidate matching areas. First, Gabor filtering and gray-level co-occurrence matrix are used to extract frequency-domain and spatial-domain directional features, respectively. Meanwhile, the parameter settings of the above methods are also discussed in order to make the extracted features correctly reflect the directional matching suitability. Then, adaptive neuro-fuzzy inference system is utilized for modeling the complementary relationship between Gabor filtering and gray-level co-occurrence matrix with the purpose of playing their respective advantages in directional matching suitability analysis. Afterward, a hierarchical decision-making scheme is designed, where the first stage is to use adaptive neuro-fuzzy inference system for selecting an appropriate analysis method (Gabor filtering or gray-level co-occurrence matrix) based on the characteristics of the given candidate matching area, and the second stage is to utilize the selected method for directional matching suitability analysis. Experimental results show that the proposed scheme is effective, and the conclusions can afford credible guidance for geomagnetic matching.

A dynamic model of a twin ducted-fan vertical takeoff and landing aircraft, the Martin Jetpack, has been developed to study and improve the understanding of the flight mechanics involved with this novel aircraft concept. This article describes the flight mechanics of a twin ducted-fan aircraft and explains in detail the modeling of the forces and moments contributed by the twin ducted-fans, body aerodynamics, control surfaces, gyration, and landing gear interactions. Also, a novel model for the movement of the duct center of pressure has been developed, which allows for the complex duct pitching moment to be predicted. Employing the conventional aircraft modeling methodology, a system of ordinary differential equations that describes the behavior of the aircraft is developed. The equations are solved in real-time using MATLAB–Simulink software to simulate the response to given inputs. A comparison of the flight data with both steady-state (trimmed) and dynamic simulations shows good agreement, which validates the novel duct center of pressure model. The validated model allows the aircraft designer/engineer to efficiently evaluate the sizing of key aerodynamic features and various control methodologies to aid in the design and flying of the Martin Jetpack.

An adaptive model of the human pilot engaged in pursuit tracking tasks that was previously introduced in the literature is modified and applied to the analysis of piloted control of a realistic transport aircraft model. As described, the pilot model requires no guesswork on the part of the analyst as regards initial parameter settings. By means of computer simulation, the adaptive pilot model is shown to exhibit superior performance to its non-adaptive counterpart in a series of configuration changes associated with the vehicle model. The overall validity of the post-adaptive pilot model is assessed by examining the resulting open-loop pilot vehicle dynamics in comparison to that predicted by the crossover model of the human pilot. The pilot modeling approach is proposed as a preliminary analytical tool to be used in the assessment of robust flight control system designs subject to faults or system failures with an eye toward potential loss-of-control.

The effects of using porous aluminum particles in solid propellants were studied, with emphasis on the agglomeration phenomena. Burning strands containing either regular (as-received) or porous aluminum were photographed by a high-speed camera, and particulate combustion products were analyzed in a laser particle analyzer. Results obtained from experiments conducted in a pressure-range of 1–34 atmospheres show that porous aluminum particles produce smaller agglomerates than regular aluminum. The median diameter of agglomerates resulting from porous aluminum reached, on average, 70% of the one originating from regular aluminum. This reduction in agglomerate diameter corresponds to a substantial volume (and hence, mass) decrease of approximately 65%. It is assumed that the high-specific area of the porous aluminum particles (10–18 m²/g, similar to that of nano-Al) results in high reactivity, leading to shorter ignition time and hence to the formation of smaller agglomerates.

Numerical experiments are carried out using commercially available Navier–Stokes solver to investigate the effect of forward-facing parabolic cavity on the heat fluxes over a spherical nosed blunt body. A wide range of parabolic cavities with depths varying between 2 and 10 mm placed at the nose of sphere-cylinder with base diameter 40 mm and overall length 70 mm have been investigated. The ratio of the cavity radius at intersection with y-axis to depth of cavity (r/d) of these cavities varies from 1.5 to 2.5. All computations have been done at a freestream Mach number of 6.2 and sea level atmospheric conditions assuming air to be a thermally perfect gas. The steady-state solutions obtained through time marching solution of axisymmetric Navier–Stokes equations suggest that the total heat transfer rate, area weighted average heat flux and the peak heat fluxes to the body can be favorably reduced for shallow parabolic cavities.

To get a better understanding on the output uncertainty contributed by an individual variable as well as the correlated variables of models with dependent inputs, a method for decomposing Sobol’s first-order effect indices into uncorrelated variations and correlated variations is investigated. Instead of using Monte Carlo simulation or full tensor product-based numerical integration approaches, a new sparse grid numerical integration method is proposed for estimating Sobol’s main effect indices as well as the two decomposed sensitivity measures. Before conducting the sparse grid numerical integration-based algorithm, an orthogonal transformation is used to transform the dependent input variables and model performance function into independent space as the joint probability density function of the correlated variables cannot be written as the product of univariate density functions. An obvious advantage of the sparse grid numerical integration-based method is that it can decrease the computational cost of the conventional methods significantly while keeping the accuracy level controllable, particularly for high-dimensional problems. The proposed approach is compared with other alternative approaches through theoretical and applied numerical experiments to demonstrate its efficiency, accuracy and high-dimensional adaptivity.

Loading an aerospace and automotive seat statically through lap or body blocks is a complex and highly non-linear problem, as the key numerical challenge is to replicate the contact and slipping kinematics between seat, lap block and belt. In addition, severe element distortions and unexpected contact between parts can occur due to the large deformations involved, which result in implicit solvers struggling to find a converged solution. This paper focuses on the use of an explicit Finite Element Analysis (FEA) solver (LS-DYNA3D) for an aircraft seat subject to Certification Specifications CS25.561, although the ideas presented are equally applicable to automotive seat designers. Explicit codes are better able to overcome contact convergence issues and are often used with appropriate damping to achieve a quasi-static solution. This paper reviews the methodology presented in Part I, whereby issues relating to damping, mass and time scaling are outlined in order to overcome the high computational time step costs (Courant-Friedrichs-Lewy (CFL) condition), together with the procedural and error checks required to ensure a quasi-static response. This paper extends the methodology by considering load cases that use lap blocks, such as ‘forward 9g’ and ‘upward 3g’ certification requirements. Alternative modelling approaches to represent the loading mechanism and effect of lap block mass on solution accuracy are discussed. This paper concludes with a verification framework that outlines the quality checks on various model energies and their ratios, where the numerical results are validated against test in terms of displacements and seat kinematics. Thus, ‘Part I’ and ‘Part II’ cover all elements related with the application of an explicit dynamic integration scheme to demonstrate static seat compliance, and together, form a clear framework to assist a Computer Aided Engineering (CAE) analyst involved in applying an explicit integration scheme to solve non-linear quasi-static analyses.

The acceleration loads of projectile during launch process of two-stage light gas gun were studied by the developed computational fluid dynamics program. With the usage of LS-DYNA software, the diaphragm rupture pressure was calculated by finite element method. The influence of different waves rupture diaphragm on the maximum acceleration loads of projectile was analyzed, keeping the configurations of gun unchanged. It is found that the maximum acceleration loads can be reduced and the muzzle velocity objective can be achieved by choosing the ruptured wave appropriately and optimizing other operational parameters. Soft launch capability is provided for launching complex lifting configuration models up to hypervelocity.

This paper presents a vibration control strategy for a flexible manipulator with a collocated piezoelectric sensor/actuator pair. A hybrid vibration controller is proposed by combining the input shaping technique with auto disturbance rejection controller. The parameters of the closed-loop system can be adjusted to the known values by disturbance compensation and linear feedback using the auto disturbance rejection controller. This way, input shaper can be designed without accurate parameters of the flexible manipulator. Both simulation and experiments are conducted to validate the proposed control algorithm. The results verified the effectiveness of the proposed controller in vibration suppression of flexible manipulator.