Journal of Aerospace Engineering

In a three-dimensional arbitrary curvilinear coordinate system, a FORTRAN program was used to research a condition’s influence on an aeroengine’s two-phase reacting flow field and combustion performance. In this program, an algebraic stress model was adopted to simulate turbulence viscosity. A second-order moment eddy–break up turbulence combustion model was used to determine the chemical reaction rate. In a nonstaggered grid system, the governing equations for the gas phase were solved by the PISO algorithm, and simulation of gas-liquid coupling was achieved by using the particle-source-in-cell method. The aeroengine combustor’s performance was estimated by using an empirical-analytical design methodology. It can be observed that performance at maximum conditions is better than that at idling conditions. The obtained agreement between simulation results and rig-test data is reasonable, which indicates the reliability of the numerical program adopted in this paper. This special-purposed program can be used to predict combustor’s flow field and its performance for the optimization design of a modern aeroengine combustor.

Direct injection into a cavity-based supersonic ramjet combustor has been recently considered to be an effective way of stabilizing the flame over a wide range of operating conditions. In the present study, two-dimensional (2D) supersonic combustor with direct fuel injection from the cavity floor with various fuel injection angles is simulated for both nonreacting and reacting conditions using Fluent software. The various complex flow features like upstream shocks, compression waves, and Mach reflection are formed in the flow field. Low velocity recirculation regions are observed in the cavity, which assists in flame holding. Heat release was observed to have significant effects on various flow structures formed both in the main stream and cavity regions. In nonreacting conditions, injection at 135° shows maximum pressure losses followed by injection angle of 120°. However, the 120° injection angle case exhibits better mixing over other cases. The present study indicates that any change in injection angle does not have any significant effect on the air entrainment. In reacting conditions, the 120° injection angle shows highest combustion efficiency because of better mixing properties exhibited by this location.

This paper presents a high-performance and simple guidance method for a short-range missile. Missile control systems are designed along the estimated flight envelope of the missile based on coefficient diagram method. The linear time invariant (LTI) models are obtained by linearizing missile nonlinear model at specific operating points. The governing equations of motion are accurately modeled by taking into account the effect of the mass rate and the center of gravity shift. A novel algebraic approach called coefficient diagram method (CDM) is used to design normal acceleration and roll angle control systems. The proportional navigation law is considered as the guidance law. One case of engagement scenarios is studied and the missile performance is evaluated. The robustness of the proposed controller is tested against parameter uncertainties and wind disturbance.

Semiglobally input-to-state stable (ISS) control law is derived for flexible spacecraft attitude maneuvers in the presence of parameter uncertainties, external disturbances, and actuator fault. As a stepping stone, a nonlinear-proportional-plus-derivative (NPD) type controller plus a sign function is developed for the fault-free system. A sufficient condition under which this NPD type control law can render the system semiglobally input-to-state stable is provided such that closed-loop system is robust with respect to any disturbance within a quantifiable restriction on amplitude, as well as the set of initial conditions, if the control gains are designed appropriately. Based on the online estimation of eventual actuator faults, this controller is redesigned for the attitude control system with actuator failures by adding a new control term to the original to reduce the actuator fault effect on the system without the need for a fault detection and isolation mechanism. In addition to detailed derivations of the new controllers design and the associated stability and attitude convergence proofs, simulation studies have been conducted to validate the design, and the results are presented to highlight the ensuring closed-loop performance benefits when compared with the conventional control schemes.

In a typical application of Kalman filtering, the filter receives the output observation of the system and the control inputs to the system and produces an estimate of the state of the system. In an air combat game, it is unreasonable to assume that direct information about the enemy’s inputs are available. Hence, a straightforward application of filtering does not work. In this paper, two different approaches are presented in estimating the states of friendly as well as enemy forces based on the output observation and the friendly control inputs when the enemy inputs are not available. Stochastic simulations are carried out in the context of a game-theoretic feedback controller and compare their performance under noise. The two methods are the Kalman filter and the unknown input-decoupling observer. The Kalman filter treats the enemy inputs as part of the extended state and obtains an estimate of both the state of the two forces and the input of the enemy. Because the filter was originally designed for a linear time-invariant system, an extension of the filter is presented to a nonlinear time-variant system. Unlike the Kalman filter, the input-decoupling observer does not need to estimate the unknown inputs. Rather, the observer is designed to decouple the unknown inputs from the estimate of the state so that the state estimate is insensitive to the enemy inputs.

In this paper, a multioutput trajectory control of a reentry vehicle is presented. Adaptive fuzzy sliding mode control with an adjustable structure of the membership function is employed. The main goal of the adjustable structure is that the span of the membership function covers all the variations of the fuzzy inputs. The fuzzy inputs include the altitude, velocity, and flight path angle of the vehicle. The sliding surface is designed in two steps. At the first step, the error of the flight path angle is solely employed. Then, to improve the performance, the altitude and velocity of the reentry are added to sliding surface. The stability of the closed loop system is proven on the basis of the Lyapunov approach. To consider performance of the closed loop system, some simulations are done. The results show outperformance of the proposed method.

Aeroelastic simulations that are computed by coupling separate computational fluid dynamics (CFD) and computational structural dynamics (CSD) codes will include not only numerical errors in both the CFD and CSD methods, but also errors associated with the data exchange among the codes of the aerodynamic loads and moments and structural deflections. Whereas the numerical errors associated with the data interface have been investigated and quantified for fixed-wing applications, their effect has not been quantified for rotational applications, such as rotors and wind turbines, in which rigid body degrees of freedom cannot be separated from flexible motion of the blades. This paper provides a systematic analysis of the numerical errors introduced by the data interface for high-aspect ratio blades undergoing multibody dynamic motion including rotation. It is demonstrated that the failure to consider conservation of work or energy during the CFD-CSD data exchange can lead to a shift in the mean loading as observed for fixed-wing applications and an alteration of the controls necessary to achieve rotor trim affecting the rotor performance prediction. Various techniques to obtain minimal data transfer errors on a given grid are examined. Data transfer on unstructured CFD meshes appear to be more sensitive than the process on structured CFD meshes.

A light cargo unmanned air vehicle (UAV) was designed, constructed, and tested in flight. This UAV was designed and built according to the specifications of the Air Cargo Challenge 2009 Design, Build & Fly European student competition. The basic aerodynamic and stability analysis that was used in the preliminary phase of the light UAV design is presented first. Flight stability analysis based on the linearized theory was performed next. Preliminary aerodynamic analysis based on Navier-Stokes solutions for wing and wing-body configurations during the design phase was carried out. Further aerodynamic analysis for the full configuration after the design was completed was carried out to evaluate the performance during the flight envelope. The findings of this analysis could be utilized to further improve the aerodynamics of the existing design and enhance stability and performance characteristics of the light cargo UAV. The conceptual design was constructed and successfully tested.

High-payload platforms flying in extreme-ground-effect represent a promising concept for high-speed amphibious transportation. Their dynamics is governed, to a large extent, by unsteady airflow under the platform. A quasi-one-dimensional model for the underplatform channel-flow is applied to predict heave and pitch motions of a vehicle in the presence of external disturbances, such as nonuniformities on the underlying ground-surface. It is found that a complete nonlinear theory should be used instead of a linear model to accurately predict motions when ground nonuniformities are relatively short. A flat platform with a flap moving in extreme-ground-effect is intrinsically unstable. An aft-wing out of ground-effect can be added to the system to achieve stability. Movable control surfaces, such as a platform flap, can be also modeled by the nonlinear theory. It is demonstrated that an oscillating flap can reduce heave and pitch amplitudes in flight over a wavy ground-surface.

An airport-based avian radar system is an effective technical means for bird-aircraft strike hazard (BASH) avoidance. Three existing systems are surveyed, including Merlin Avian Radar System, Accipiter Avian Radar System, and Beihang Experimental Avian Radar System. Two critical technologies for Beihang experimental system are introduced: flying bird target detecting and tracking algorithm and BASH assessment. The algorithm consists of five steps: background subtraction, clutter suppression, measurement extraction, multitarget tracking, and data overlay. The multitarget tracking algorithm is tested against simulated data. The overall approach is applied on a set of horizontal scanning plane position indicator (PPI) radar images. A BASH assessment model is established by an analytic hierarchy process (AHP), considering two factors of the current states of flying birds and aircraft. Using the BASH assessment model, an application example at Nanyang Airport is described, in detail. Finally, some development trend of critical technologies for airport-based avian radar system is given.

In this paper, an improved self-learning simulation (SelfSim) method is proposed for the inverse extraction of nonuniform, inelastic, and nonlinear material behavior under cyclic loadings. The SelfSim has been used to inversely extract local inelastic and nonlinear behavior of materials using limited global boundary responses. However, the SelfSim with conventional artificial neural network (ANN) models needs ad hoc data processing that frequently interrupts SelfSim training even in training monotonic constitutive behavior. To overcome this problem, an improved SelfSim with a new ANN-based hysteretic model is proposed. In addition, the ANN material model is implemented with considerations of large volume changes and geometric nonlinearity. The new SelfSim shows superior performances in the inverse modeling of complex material behavior under “multiaxial” and “cyclic” stress states. Two simulated numerical tests using a laminated rubber bearing with reinforcing steel shims are used to demonstrate the proposed SelfSim performance. A detailed comparison of the SelfSim with conventional ANN model is also presented. Finally, the improved SelfSim is experimentally verified by extracting nonuniform, nonlinear, and inelastic behavior of metallic material (low carbon SAE 1006 specimen) under cyclic loading. It shows very promising performances for the inverse material characterization of various engineering materials.

The main goal of the work described in this paper was to set up a procedure for modeling a thermal protection system for a hypersonic reentry vehicle by extending earlier work done by the authors. A multiphysics framework has been setup for the simulation of hypersonic reentry vehicles using commercial codes FLUENT and LS-DYNA with user-defined programming. The computational fluid dynamics code (FLUENT) and the material thermal and structural response code (LS-DYNA) are loosely coupled to achieve the solution. The surface recession resulting from ablation in the reentry vehicle is simulated by implementing a mesh movement algorithm in LS-DYNA. The vehicle considered in the calculation is an axisymmetric vehicle flying at zero degree angle of attack.

This work extends advanced beam models to carry out a more accurate free-vibration analysis of conventional (straight, or with sweep/dihedral angles) and joined wings. The beam models are obtained by assuming higher-order (up to fourth) expansions for the unknown displacement variables over the cross-section. Higher-order terms permit bending/torsion modes to be coupled and capture any other vibration modes that require in-plane and warping deformation of the beam sections to be detected. Classical beam analyses, based on the Euler-Bernoulli and on Timoshenko beam theories, are obtained as particular cases. Numerical solutions are obtained by using the finite element (FE) method, which permits various boundary conditions and different wing/section geometries to be handled with ease. A comparison with other shell/solid FE solutions is given to examine the beam model. The capability of the beam model to detect bending, torsion, mixed and other vibration modes is shown by considering conventional and joined wings with different beam axis geometries as well as with various sections (compact, plate-type, thin-walled airfoil-type). The accuracy and the limitations of classical beam theories have been highlighted for a number of problems. It has been concluded that the proposed beam model could lead to quasi-three-dimensional dynamic responses of classical and nonclassical beam geometries. It provides better results than classical beam approaches, and it is much more computationally efficient than shell/solid modeling approaches.

Quantifying the strain rate–sensitive mechanical properties of structural materials is an important area of research in the field of solid mechanics. Property evaluation is typically accomplished using dynamic tests that involve rapid loading or impact of specimens. These tests generate inertial forces and wave propagation, which make it difficult to accurately record the material response to a loading condition at an equivalent location. Furthermore, dynamic impact tests typically generate high strain rates (in excess of ${10}^3{\mathrm{s}}^{-1}$) and an experimental method for generating rates of strain in the intermediate strain rate regime that is simple and reliable is still lacking. This research develops and investigates an experimental technique for generating tensile plastic strain rates up to ${10}^2{\mathrm{s}}^{-1}$ in ductile metals. The technique relies on the impact from a vertical drop weight machine capable of delivering suitable impact velocity and energy to globally deform a slotted beam specimen. At impact, a state of predominantly plastic uniaxial tensile stress is created in the ligament beneath the slot. The ligament is instrumented with an electrical resistance strain gauge, and the strain history is measured and stored in a digital oscilloscope. Experiments are conducted on commercially pure (CP) titanium, 2024-T3 aluminum, and 1018 steel samples at four distinct impact velocities, and collected data illustrate the capability to create predominantly uniaxial tensile strain rates up to ${10}^2{\mathrm{s}}^{-1}$. An analysis of the material response identifies where plastic flow initiates. Furthermore, a numerical analysis of the impact event is conducted where the Johnson-Cook constitutive equation is assumed to reflect the material behavior and published parameters are utilized to illustrate good agreement between experimental strain data and the numerical model.

The mechanical response of epoxy resin Epon E 863 has been studied in tension, compression, and flexure. The epoxy resins have been tested at different strain rates ranging from $5.9\times {10}^{-5}$ to $0.03{\mathrm{s}}^{-1}$. Two types of dog-bone geometries have been used in the tension tests. Small sized cubic, prismatic, and cylindrical samples were used in compression tests. Beams with quarter deep notches or grooves were tested at their midpoints in flexural tests. Strains were measured by using a digital image correlation technique, extensometer, strain gages, and actuator. Observation of sample geometry during tension tests at constant elongation rate shows necking and crazing in Epon E 863. Cubic, prismatic, and cylindrical compression samples undergo a stress drop at yield, but only cubic samples experience strain hardening before failure. Characteristic points of tensile and compressive stress strain relation and load deflection curve in flexure, such as proportional elastic limit stress (PEL), ultimate tensile strength (UTS), compressive yield strength (CYS), and failure strain were measured and their variations with strain rate studied.

The robust vibration control and dynamic response analysis of an advanced aircraft wing structure are addressed in this paper. The wing structure is modeled as a thin-walled beam with fiber-reinforced composite materials featuring a circumferentially asymmetric stiffness (CAS) configuration that completely splits into two coupled groups of motion. The time-dependent external loads are considered with unsteady aerodynamics in an incompressible flow field. To demonstrate the robust characteristics of the sliding mode methods, both the conventional linear quadratic Gaussian (LQG) regulator and the proposed method are compared with respect to model uncertainty and external excitation. The numerical results demonstrate that the sliding mode control (SMC) on the basis of the sliding mode observer (SMO) is an efficient way to control the unstable dynamic response subjected to various external loads associated with model uncertainty.