Journal of Aerospace Engineering

Any configuration consisting of flat plate attached to a ramp experiences shock wave boundary layer interaction (SWBLI) at supersonic or hypersonic Mach numbers. This interaction can lead to flow separation at the ramp foot. In that case, its extent depends on the leading-edge radius of the plate. At a specific radius (inversion radius), the separation bubble size attains maximum value, and at particular leading-edge bluntness (equivalent radius), the length of the separation bubble becomes equal to that of the sharp leading-edge case. Values of these critical radii need to be evaluated for complete understanding of dynamics of SWBLI. Therefore, numerical simulations are carried out to estimate these values with change in freestream Mach number, wall temperature, and freestream stagnation enthalpy. Thicknesses of the high-entropy layer and boundary layer, along with wall pressure distribution, are considered for analyzing the influence of these parameters on the estimation of critical radii. The magnitudes of both the radii have been observed to increase with a decrease in Mach number and stagnation enthalpy. Freestream static and total temperatures are found to have dimensional dependence on the flow field for moderate leading-edge radii instead of their ratio. Therefore, present studies are useful in understanding and accessing the effect of freestream conditions and wall temperature on the values of critical radii.

This paper proposes a new singularity direction escape steering law for control moment gyros (CMG) using singular value decomposition (SVD). The proposed steering law differs from the classical singular direction avoidance (SDA) steering law in two respects. First, torque errors are introduced in the direction of the smallest two singular values through modification of the diagonal matrix. This is done by tuning parameters related to the magnitude of torque error so that both singularity avoidance and escape performance can be achieved. Second, instead of using a conventional singularity measure, new output torque metrics of the CMG cluster along the torque command direction are introduced and adopted in this steering law, so as to reduce unnecessary torque error for certain near-singular circumstances. Several numerical simulations are provided at last, and comparisons with existing steering laws are presented to demonstrate the superior performance of the proposed method.

A generalized predictive control with adaptive fuzzy estimator (GPC-AFE) is proposed for spacecraft. Model uncertainties and mismatch disturbances are considered in the process of controller design. For the attitude subsystem, a predictive control based on an adaptive fuzzy estimator is designed to deal with the unknown disturbances. For the angular velocity subsystem, the actual control input is designed by using the generalized predictive control. The derivative of virtual control and unknown disturbances in the actual control are also estimated by using an adaptive fuzzy estimator. With the proposed generalized predictive control, the tracking error of the attitude converges to a bounded neighborhood of the reference attitude. Theoretical results are illustrated by numerical simulation.

Generalized constitutive models, which can be used in explicit finite-element analysis, are being developed to accurately model composite systems under impact conditions. These models require extensive characterization of the composite material as input—the more sophisticated the model, the larger the amount of required input data. The authors discuss details of the characterization data required for a generalized three-dimensional orthotropic elastoplastic material model that is under development to simulate the impact response of composite materials. The material response data are generated using virtual testing methods in an analytical approach based on the generalized method of cells and a finite-element system. These methods not only complement actual test data by providing an additional layer of verification for experimentally generated stress-strain curves; they also fill in gaps caused by lack of appropriate experimental data (due either to the required experiments not being carried out or to the experiments being too complex to carry out). Verification and validation test cases demonstrate the capabilities of the developed system and show how a general virtual testing framework can be used for characterizing orthotropic composites.

The locations of lumped masses, e.g., engines, missiles, fuel tanks, etc., can alter the dynamic characteristics of the airframes, thereby affect the critical flutter speed and gust response of the wing. Usually the practical location of the lumped mass is uncertain due to errors in processing and manufacturing; therefore, frequency interval optimization of the wing under uncertain locations of lumped masses is developed in this paper. The uncertain locations of the lumped masses are described by interval variables; the optimization problem with interval variables can be transformed into an equivalent deterministic problem by the first-order Taylor series of the objective and constraint function, and then the optimal locations of the lumped masses can be obtained by the sequential quadratic programming algorithm. A straight wing modeled by beam elements, with uncertain locations of lumped masses, was studied to demonstrate this method, and the results show that the bounds of the optimal results can be narrowed by improving the accuracy of the locations of the lumped masses. The example of a swept wing is given in this paper to illustrate that the frequency interval optimization procedure can be done by integrating commercial software so this method can easily be applied to solve the complex problem in practice. The results show that interval optimization can give the bounds of these optimal frequencies, and increase the frequency difference between the second bending mode and the first torsion mode, which can increase critical flutter speed.

When a rocket lands on the Moon, the exhaust plume causes the lunar regolith (soil) to be ejected in all directions. Such high-speed ejection poses a threat to nearby equipment and structures. As a step toward better understanding the ejection process and mitigating the corresponding destruction, the discrete element method (DEM) is used to investigate the erosion flux of soil with a (continuous) lognormal distribution of particle size. The results show that the erosion flux increases relative to a monodisperse (single particle size) case with the same Sauter mean diameter due to an increase in the erosion flux of large particles and a reduction in that of small particles. The physical underpinnings of this counterintuitive behavior involve the interaction between small and large particles, particularly through collisions and static packing. Namely, the faster-moving smaller particles preferentially transfer momentum via collision to larger particles from a lower vertical position, causing an increase in large-particle flux and vice versa for small particles. Polydisperse mixtures also pack tighter than monodisperse mixtures, thereby increasing the erosion flux of both species by increasing the bulk density of the soil and the eroding material. Finally, to facilitate the use of kinetic-theory-based continuum models for this application, the discretization of continuous distributions is explored (e.g., representing a Gaussian distribution with a discrete approximation, e.g., particles of five different sizes). Two techniques are considered: the method of matching moments and a new method known as volumetric discretization. For the parameter space studied here, the volumetric discretization performs better when comparing the erosion flux of the continuous and discretized distributions.

Vortex generator jets (VGJs) have been used to suppress the flow separations in bowed compressor cascades at different incidences via numerical calculations. The results show that with momentum injection and streamwise vortex produced by VGJ, the total losses in the bowed blades effectively decrease whereas the static pressure ratios increase. For the positively bowed blade, at $-5°$ incidence, with slight flow separation, the mixing loss caused by VGJ is dominant and the loss increases by 1%; with an incidence ranging from 0° to $+5°$, the regions of the flow separations all decrease, and the total losses decrease by 2.7 and 5.4%, respectively; with the incidence increasing to $+7°$, the separation type converts from closed to open, and the loss decreases by 6.8%; when the incidence is $+10°$, one concentrated shedding vortex disappears, and the loss decreases by 9.27%. For the negatively bowed one, as the incidence increases from $-5°$ to $+7°$, the flow separations are all effectively suppressed when all the separation types are closed, and the losses decrease by 2.9, 9.1, 17.5, and 19.75%, respectively. When the incidence increases to $+10°$, however, because of the huge range of flow separation, there is no improvement in the flow field, but the total loss increases by 1.37%.

The paper develops a noncertainty-equivalence adaptive (NCEA) spacecraft formation control system, using filtered signals, based on the immersion and invariance methodology. It is assumed that a target spacecraft is in an elliptic orbit and a follower satellite is moving around it. It is also assumed that the mass of the follower satellite is not known, and its dynamics include time-varying periodic as well as random disturbance forces. The objective is to design an adaptive control system so that the follower spacecraft remains in a specified formation with respect to the target spacecraft. First, based on the immersion and invariance theory, a control system—consisting of an adaptive stabilizing law and a parameter identifier—is designed for the relative position control of the follower satellite, perturbed by time-varying periodic forces. The control system is synthesized using filtered signals. Unlike traditional certainty-equivalence adaptive laws, the parameter estimates include certain nonlinear algebraic functions, besides signals obtained by integral action. Based on the Lyapunov approach, it is shown that all the signals in the closed-loop system are bounded, and that the relative position trajectory error asymptotically converges to zero. A special feature of the control systems is that the trajectories of the closed-loop system eventually evolve on an attractive manifold in an extended state space. Furthermore, the parameter identifier has strong stability properties. Although the NCEA system is robust to disturbance inputs, $\sigma$ modification is introduced in the update law for regulating the residual set, and global uniform ultimate boundedness of the trajectories is established. Simulation results are presented which show that the designed control system achieves precise formation control, despite parameter uncertainty and time-varying periodic and random disturbance forces.

A time efficient robust proportional, integral, differential (PID) plus controller for satellite attitude stabilization control is proposed in this paper. A time efficient sliding mode with two-stage structure is designed to improve the system convergence rate and a standard PID controller is modified to ensure the system state reaches the sliding mode. The controller in this paper does not need an accurate system inertia matrix and is robust to unknown disturbance torque with a norm upper bound. The angular velocity and control torque constraint to ensure the system states do not exceed their upper bound is discussed. A novel method to estimate the norm of angular velocity is proposed. The globally asymptotic stability for all conditions considered in this paper is proved, and the sliding mode state is also proved to have global asymptotic stability. The performance of the controller is demonstrated by simulation results.

The results of an experimental and analytical study on aircraft radial tire blowout are presented in this paper. A new experimental method was introduced to study the flow field of tire blowout, including the design of an airfoil-shaped support rod and the arrangement of transducers. In this study, several radial tires with the same blowout pattern were tested and variations in blast pressures in time at all test points were presented. Based on the test data, an analytical model accurately describing the peak pressure distribution in the flow field of radial tire blowout was derived. Using the European Aviation Safety Agency (EASA) model and approximation results, the peak pressure distributions at angles of 0, 15, 30, and 45 degrees from the blowout centerline and at distances of 100, 200, 300, 400, and 500 mm to the tire face were investigated. It was shown that pressure distribution is dependent on the angle from the centerline and the distance to the tire face. In addition, variations in equivalent static pressure with angle and distance were analyzed. The transient process of radial tire blowout was presented through high-speed video data, and the fracture mode of radial tires was summarized. Finally, the potential effects of tire blowout on nearby equipment and personnel were studied and safety precautions suggested.

This work focuses on defining loads and fatigue life estimation of helicopter tail rotor blades. The original metal construction of these blades was replaced with a new construction of blades made from composite materials. The subject of this investigation is the precise determination of loads of helicopter composite tail rotor blades together with experimental verification of their fatigue strength. In this research to determine loads of tail rotor blades computational fluid dynamics (CFD) numerical simulations are used. To study the fatigue behavior of composite blades, the blades are here tested under static and fatigue load spectra. For determination of aerodynamic loads of tail rotor blades in this investigation, two computation models are used. In first model, an isolated tail rotor blade is considered. In the second computation model, the complete helicopter is modeled, including fuselage together with main and tail rotor blades. The tail rotor hub is not included in these computational models. To verify the strength of the composite structure of the tail rotor blade, it is tested under the static load and fatigue load spectra. The experimental fatigue test results under the load spectrum are included.

In the reentry trajectory optimization for unpowered hypersonic glide vehicles (HGVs), the initial reentry conditions (IRCs), which are crucial for flight performance, are partly flexible instead of being specified. Aiming at simultaneously determining the optimal IRCs and optimal control policy, a hybrid optimization strategy with an evaluate-during-optimize framework is proposed in this paper. The IRCs are considered as additional independent design variables, and an evaluation process with multiple performance indicators is established to determine the best IRCs. With the analytic hierarchy process (AHP)-based multicriteria decision-making (MCDM) method, a synthesized scalar objective function is set up. Adopting this objective function, feedback information from the evaluation process can be fully included in the optimization process, resulting in a unified optimization framework. The Chebyshev pseudospectral method (CPM) is then introduced as a direct method for solving the resulting trajectory optimization problem. Numerical simulation results for a generic HGV show that the proposed technique is efficient and straightforward. Moreover, this strategy has strong universality for solving optimal control problems with flexible initial conditions.

Large eddy simulations (LES) are employed to investigate instabilities in the combustion process of a model scramjet combustor at the German Aerospace Center (DLR). Dynamic mode decomposition (DMD) is applied to analyze the numerical results. The DMD spectrum of the hydroxyl (OH) field is analyzed, and primary peaks are observed at 6375, 9763, 12063, and 15745 Hz, corresponding to ${\mathsf{S}}_t=0.0524$, 0.0802, 0.099, and 0.129, respectively. The DMD mode corresponding to ${\mathsf{S}}_t=0.099$ exhibits antisymmetric patterns in the wake of the strut, with the DMD mode corresponding to ${\mathsf{S}}_t=0.129$ displaying similar structures. The DMD spectrum of the hydroperoxyl (${\mathrm{HO}}_2$) field is also investigated. Antisymmetric patterns initiating from both trailing edges of the strut are also extracted at nearly the same Strouhal numbers. DMD analyses are also performed on the pressure and velocity fields. The dominant frequencies in the pressure field are lower than those of the OH and ${\mathrm{HO}}_2$ fields, indicating that the dominant pressure oscillation is unlikely to couple with the strongest unsteadiness in the OH and ${\mathrm{HO}}_2$ fields. The streamwise component of the velocity DMD mode attributed to the same Strouhal number exhibits an antisymmetric pattern similar to the most unstable OH DMD modes. Wake instability is believed to be the main cause of the dominant oscillations in the OH field. Moreover, the unsteady characteristics of the ${\mathrm{HO}}_2$ field are also related to the wake instability.

Spray characteristics of gas-centered swirl coaxial injectors, which are commonly used in staged combustion cycle engines, are investigated numerically and experimentally. The spray angle and the spreading angle are the quantitative parameters selected for experimental and numerical studies, respectively. Various gas/liquid Reynolds numbers are adopted for the operating conditions. For a fixed gas Reynolds number, the spreading angle increases with the liquid Reynolds number. Through the comparative study of the spray angle from water/air experiments, the spreading angle is analyzed for different momentum flux ratios (MFRs). The comparison of the spray and spreading angles for different MFRs shows good agreement. Modeling and study of gas–gas flow is a useful approach to understanding spray characteristics of gas–liquid flow.

This paper studies the unsteady laminar magnetohydrodynamic (MHD) mixed convection flow of an incompressible viscous non-Newtonian Casson fluid in the stagnation region of an impulsively rotating sphere. Nonsimilar solutions and the fourth order Runge-Kutta method are used to solve the partial differential equations governing the problem. The obtained results are presented in terms of the velocity profiles in the $x$ and $y$ directions, the shear stresses profiles in the $x$ and $y$ directions and the profiles of heat transfer rate. The outcomes of the paper indicated that as the Casson parameter increases, the velocity profiles in the $x$ direction increase, whereas it reduces the velocity profiles in the $y$ direction and temperature distributions. Additionally, the increase in the shear stresses in both the $x$ and $y$ directions can be obtained by increasing either the magnetic field parameter or buoyancy parameter.

The behavior of nonlinear oscillations of isotropic rectangular plate fluttering in a supersonic gas flow is examined. The study was conducted taking into account both types of nonlinearity: wind (quadratic and cubic) and geometric (cubic). It is established that because of aerodynamic nonlinearity (especially its nonsymmetric quadratic part), the relationship $A(\nu )$ (where $A$ is the amplitude of nonlinear oscillations and $\nu$ is the parameter characterizing the value of the flowing stream) is a two-value function at certain intervals of the speed $\nu$. This fact is illustrated in the figures, which are plotted in the text, in the form of two branches. The lower branches of these figures, in all probability, are unstable. The unstable branches are separated via the gravitational field of two adjacent sustainable solutions. Thus, the perturbation magnitude, which is required in order to transfer the system from one stable branch to another, can easily be found. The existence of certain intervals of $\nu$ is shown, where limit cycle oscillations (LCOs) cannot be excited in both precritical speeds and in the postcritical stage.

A nonlinear optimal control technique is utilized to address coupled relative rotational and translational motion in spacecraft rendezvous. A new formulation of relative rotational dynamic with momentum exchange devices is developed. For a complete description of the relative motion dynamic in the final proximity phase, a kinematically coupled model of relative position and attitude is derived. For the sake of safety of the maneuvers and achieving smooth paths, which are critical issues during proximity, accurate motion control is needed; as a result, the state-dependent Riccati equation control technique is used to address far-range rendezvous and final proximity to a target in an elliptical orbit under small bounded perturbations and parameter uncertainty. The resulting state-dependent algebraic Riccati equation is solved using eigenvectors of a Hamiltonian matrix to find nonlinear optimal control gains. The proposed direct and numerically stable Hamiltonian method decreases computational time and cost. Numerical simulation results show that with this control method acceptable robustness and stability can be achieved.

A recently emerged driving requirement for microsatellites is a precise and fast attitude maneuvering capability. Agility enhances the operational efficiency of missions requiring a high level of retargeting ability, such as space station and satellite on-orbit servicing, formation flying, Earth monitoring, and space-based space debris observation and tracking. Three-axis stabilized microsatellites are mainly equipped with momentum wheels, barely meeting agility requirements. In this paper, the design and manufacturing of a cluster of four single-gimbal variable-speed control-moment gyros (SGCMGs), suitable for attitude control with high-agility capability of a 10–30 kg class microsatellite, is presented. The flywheel design is based on a brushless motor out-runner configuration, in which the motor magnets, embedded in the rotor, provide for the wheel inertia. This configuration has the advantage of providing a high inertia while minimizing cogging torque at low speed. The elimination of the wheel dummy mass provides a significant improvement in the system mass efficiency. In addition, this configuration allows using commercial off-the-shelf parts, eliminating the wheel-balancing procedure with evident jitter limitation and economic and development time benefits. A relevant feature is the electronics architecture based on a field-programmable gate array (FPGA) integrated circuit, chosen to improve power efficiency and provide a high level of reliability for the system. This can be assured by proper time and space redundancy for each low-level logic block of the design. The system performance in terms of provided torque and power consumption has been experimentally measured by using a ground-testing prototype mounted on a test bed, and the results are reported. The results confirm the effectiveness of the proposed actuator as an efficient solution for the attitude control of microsatellites, in terms of mass, volume, and power constraints.

The ability to fold and deploy lightweight composites without damage makes them attractive for aerospace applications. However, one of the challenges faced with deployable composites is their high stiffness, which results in a relatively high deployment rate. It has been hypothesized that by exploiting the time-dependent viscoelastic response of composites, the deployment process could be controlled. To investigate this hypothesis, the effect of matrix viscoelasticity on energy dissipation of a three-layer carbon fiber–reinforced polymer (CFRP) composite laminate, known as a composite tape spring, was examined during the stowage state. A time-dependent implicit finite-element model was generated and implemented to simulate the viscoelastic behavior of the orthotropic laminated CFRP composite tape spring. The implemented material model was verified against data from the literature, validated experimentally, and then used to investigate the significance of matrix stress relaxation on energy dissipation of the three-layer CFRP composite tape spring used in aerospace applications.

The development of new rocket engine nozzles for launch vehicles encounters a challenging design problem: to meet nozzle performance for high altitude, the nozzle expansion ratios are designed by high value; however, these would lead to overexpanded flow conditions at ground. These conditions in turn generate unsteady internal flow separation. The resulting asymmetries result in side loads, which can potentially damage not only the nozzle but also the entire launch system. The occurrence of excessive side loads in nozzles is one of the most important issues to consider in designing efficient, reusable, and robust launch vehicles. In this work, a fully coupled method using Navier–Stokes simulations is used to numerically investigate the aeroelastic stability for the J-2S rocket nozzle by varying either the material properties or the thickness of nozzle walls. This fully coupled method consists of the following: (1) a flow solver to simulate the flow field, (2) a structural solver to compute the structural dynamic response, (3) a computational mesh dynamics solver to accomplish the deformation of the fluid dynamics grid, and (4) a coupling technique to swap forces and displacements across the fluid–solid interface. It is found that the wall material properties and the wall thickness have tremendous effects on the aeroelastic behavior of rocket nozzles. Thus, the interaction between the shock-encompassed flow inside the rocket nozzle and the rocket nozzle structure has to be considered for the design of rocket engines.

This paper addresses the problem of precise in-flight simulation of all dynamic engine operational modes, considering the specific features of the control in these modes. One of the dynamic processes that are most complex to simulate is “cold stabilization”—that is, acceleration with incomplete heating of engine elements. The principal peculiarity of this process is associated with a greater clearance between the blade tips and the casing facing them, which results in lower efficiency of the compressor or turbine. Hence, engine control programs that are developed for a heated engine tend to change the engine static performance (e.g., greater fuel consumption and higher gas turbine inlet temperature) and make the transients slower when applied to cold stabilization. The development of active clearance control systems makes evident the need for fast mathematical models that consider the clearance change at different dynamic operational conditions. Clearance is determined by the radial sizes of the disk, blades, and casing. Therefore, the main factors affecting it are thermal strains and strains generated by mechanical loads that change these sizes. Numerous studies have addressed thermal straining simulation, but only a few have considered straining due to mechanical loads. However, even these works do not consider the influence of disk temperature on the strains induced by mechanical loads. Indeed, these strains depend on temperature because the mechanical properties of the disk, especially the elasticity modulus, are temperature dependent. This paper describes a method to estimate the radial elongation of an engine heated element (HE) caused by heating and mechanical load, when the action of the latter takes into consideration the element temperature. The proposed method was experimentally validated for the radial clearance dynamics of a high-pressure turbine disk. Finite element analysis and real data proved that the accuracy of the method is high enough. Moreover, the method is relatively simple and can be integrated into an engine dynamic model; moreover, it does not significantly increase total time of the execution of the model program. In this way, this clearance evaluation method can be recommended for practical use.

Trajectory optimization of hypersonic vehicles has become a hotspot because of its advantages in flight speed and range. This task is challenging because of the strong nonlinear couplings among the aerodynamics, propulsion, and flight state of hypersonic vehicles. In this paper, a novel chicken swarm–based method is proposed to solve the ascent trajectory optimization problem. Firstly, the ascent trajectory optimization problem is formulated into the optimal control problem, and the cost function that ensures the minimum fuel consumption, which is subjected to various constraints such as dynamic pressure, load factor, and aerodynamic heating, is proposed. Then the principles of the chicken-swarm optimization (CSO) algorithm are explained, and the CSO-based trajectory optimization method is proposed. With the method, the control variables are discretized at a set of Chebyshev collocation points, and these points are the variables to be optimized in the CSO algorithm. Also, a segmented strategy is introduced to treat the constraints discriminately in the penalty terms of the cost function according to the flight state of the hypersonic vehicle. Based on such a strategy, the process of trajectory optimization with the CSO algorithm is depicted. A series of comparative experiments were conducted to investigate the feasibility and superiority of the proposed method. Furthermore, more experimental results are presented to discuss the influence of hierarchical order and collocation point selection on the performance of the CSO-based method.

Recently, multivariate simplex B-splines (MVSB) function approximators have been investigated with the aim of providing accurate global aerodynamic models for use in adaptive flight control systems. In this paper, a new approach for constructing multivariate spline models is presented in the form of the tensor-product MVSB (TP-MVSB) that consists of tensor products of ordinary MVSB. The key advantage of this new approach is that it provides more flexibility in the definition of the spline model structure than the standard multivariate spline approach. This flexibility allows the user to include a priori (expert) knowledge of the system in the definition of the spline model structure leading to more efficient and physically meaningful models. The TP-MVSB maintains the desirable properties of the MVSB in the sense that the global B-form regression vector is normalized, each basis polynomial is guaranteed to be well-conditioned numerically, and differentiability is maintained along each input dimension. The new approach is validated using data obtained from a nonlinear F-16 model. Simulation results show that the new approach can achieve a higher level of approximation accuracy using fewer parameters when modeling the aerodynamic moment coefficients, and in addition can provide accurate estimations of the control effectiveness matrix in cases where the system is affine in the inputs.

In this paper, the transient behavior of an internally damaged laminated composite plate and shell structure under the influence of different mechanical loading types and constraint conditions has been analyzed numerically. For the numerical purpose, two well-known higher-order displacement kinematics are used to model the doubly curved shell panel in association with the finite-element steps. In addition, the internal delamination is modeled with the help of two sublaminate approaches including the intermittent continuity condition to obtain the necessary solutions. Further, the domain has been discretized with the assistance of a biquadratic nine-noded quadrilateral element. The panel motion equation is derived by integrating the total Lagrangian expression and solved to evaluate the time-dependent responses via an in-house computer code in association with Newmark’s direct integration scheme. The stability of the found numerical solutions are checked through a convergence test and compared with established benchmark solutions. The performance of the developed numerical models are established by comparing the results with the subsequent experiments. Finally, the effect of internal debonding (size, position, and location) and other design parameters on the time-dependent deflections of the delaminated composite panel are examined including the geometries (spherical, cylindrical, elliptical, and hyperboloid) and discussed in detail.

Unmanned platforms with a conventional fixed-wing configuration can be designed according to well-assessed aircraft sizing techniques. Conversely, the design of multirotor vehicles is still based mainly on scale-model enthusiastic experience. At present there has been a limited effort by the community in developing a systematic approach for sizing this unconventional class of flying vehicles. This paper proposes a revised version of a classical aircraft sizing methodology based on statistical data available in the literature, with the objective of preliminary sizing of an electric multirotor configuration, taking into account mission profile and a few performance requirements such as endurance or percentage of thrust.

Based on fewer uncertain parameters of aeroengine blisks in statistics, a new measurement methodology is presented, which is called the interval nonprobabilistic analysis based on the dynamic substructural collaborative finite-element method (INP-DSFEM). Firstly, on the basis of the basic theory of nonprobabilistic reliability analysis, the definition of uncertainty and its extension are proposed. The expression formula of performance function is also deduced. Then, the interval variables and interval matrix of blisks are investigated, including the reduced interval bladed model, reduced interval disk model, and comprehensive hybrid interface reduced interval blisks model; moreover, the interval finite-element dynamics equation is also built. The basic thought of this method is to compare the performance range with the structural requirement range to determine the extent of security and reliability. The system will be at failure if the performance range is beyond the requirement range; however, if the performance range is in the requirement range, the system will be safe. When the upper and lower performance boundaries are not near the structural requirement, the structural range limitation of the uncertain parameter is smaller, and the structural robustness parameter is greater. It is a convenient calculation, and the permission variation boundary of uncertain parameters can be given explicitly. Finally, the nonprobabilistic analysis of blisks is investigated by INP-DSFEM, and the uncertainty of modal and vibratory response for tuned and mistuned blisks are calculated by INP-DSFEM; compared with those for the tuned blisks, the uncertainties of the modal and vibratory response for the mistuned blisks decreased. Meanwhile, a method called interval nonprobabilistic-ellipsoidal convex model (INP-ECM) is adopted to verify the reasonability and effectiveness of INP-DSFEM. Therefore, the presented methodology in this paper is a choice for the reliability calculation of aeroengine blisks, which is as a beneficial supplement for the probabilistic analysis.

Control of a Mach 2 elliptic jet from a convergent-divergent elliptic nozzle of Aspect Ratio 3, with rectangular and triangular tabs of identical blockage, was studied experimentally at different levels of pressure gradient at the nozzle exit. At all levels of pressure gradient studied in this work, irrespective of tab geometry, the tabs along the minor axis promoted mixing and the tabs along the major axis retarded mixing. However, the level of mixing was sensitive to the tab geometry. The triangular tabs along the minor axis caused greater mixing than the rectangular tabs at all pressure gradients. The triangular tabs along the minor axis caused a maximum core length reduction of 77% at a Nozzle Pressure Ratio (NPR) 4 and a minimum of 64% at NPR 5. The maximum and minimum core length reductions caused by the rectangular tabs along the minor axis were 39%, at NPR 6 and 20%, at NPRs 5 and 8. Shadowgraphs of the jet show that the waves in the core were rendered weak by both rectangular and triangular tabs when placed along the minor axis.

The solar spectrum can be obtained by a spectrometer on a spacecraft. The velocity of a spacecraft with respect to the Sun can be calculated according to the solar frequency shift. However, small-scale activities on the solar surface can cause spectral line drift, which results in a large and varying bias in the velocity measurement. This makes it difficult for a single spacecraft to exploit the solar frequency shift. To resist the interference of solar surface activities effectively, this paper develops a solar frequency shift–based radial velocity difference measurement method to enhance the accuracy of relative navigation for formation flight. In this method, each spacecraft in a formation observes the solar frequency shift and obtains a radial velocity measurement which includes the velocity bias caused by the solar surface activities. The difference in these radial velocities is immune to the varying velocity bias. Therefore the difference is utilized as the navigation measurement in this method. However, this navigation system cannot work alone because of unobservability. To make the navigation system observable and improve the accuracy of both absolute and relative navigation for formation flight, this paper combines it with X-ray pulsar navigation and proposes a radial velocity difference/pulsar integrated navigation method with observability. This autonomous integrated navigation scheme adopts the extended Kalman filter to deal with the radial velocity difference measurement and the pulsar time of arrival (TOA), providing absolute and relative navigation information for formation flight. The simulation results demonstrate that both the absolute and relative navigation accuracies of the proposed method are higher than those of X-ray pulsar navigation, especially relative navigation.

This paper presents a numerical investigation of active aerodynamic flow control on an airfoil by dielectric barrier discharge plasma actuators. Simplified Maxwell equations for an electric field coupled with Navier-Stokes equations and turbulence models for fluid are implemented to simulate the wall jet flow induced by the body force attributable to the dielectric barrier discharge plasma actuators. The numerical model is validated with the laboratory results in the quiescent flow. The simulated maximum velocity in the quiescent environment is proportional to the applied voltage. The dielectric barrier discharge plasma actuators are also demonstrated to significantly influence the velocity profiles. The numerical model is applied to an airfoil with Reynolds number equal to ${10}^6$, and the peak-to-peak voltage varies from 5 to 18 kV. The flow separation is delayed or removed to some extent by dielectric barrier discharge plasma at high angles of attack. The dielectric barrier discharge increases the lift coefficient, reduces the drag coefficient, and finally improves the performance of the airfoil.

Two attitude coordination controllers that make the system converge to the equilibrium point in finite time are designed for spacecraft formation flying with actuator saturation constraints. The first finite-time attitude coordination controller is suitable for the situation that all states of the closed-loop system are available. To deal with the problem of failing to measure angular velocity, a novel filter is presented in the second finite-time attitude coordination controller. Because the model of spacecraft formation flying is established based on the rotation matrix, the closed-loop system does not have the drawback of unwinding. The theoretical analysis and simulation results are given to prove the validity of the proposed controllers.

An accurate and efficient gridless method is presented for calculation of unsteady flows at low and high reduced frequencies. A central gridless scheme is applied to solve unsteady flow equations in the arbitrary Lagrangian-Eulerian formulation. Taylor series least squares are used to discretize spatial derivatives at each node. Constrained weights optimization is proposed using a blend of second and fourth differences artificial dissipation terms. The proposed scheme leads to a faster gridless method with simple formulations and a reduction in flux computations. For time advancement, explicit and implicit dual-time methods are used. The segment spring analogy is applied to control/treat smoothness of the dynamic clouds and nodes movement. The capability and accuracy of the method are examined by comparing results of a multielement airfoil in steady flows and some test cases in unsteady flows with those from a finite-volume method and experimental data for low angles of attack at the prestall region. Results show a good agreement with the experimental data. Moreover, it is shown that the constraint weights optimization on the fast gridless method increases the efficiency and accuracy of the approach.

Three-dimensionality effects have been investigated for a flat plate undergoing an impulsive pitch-up motion in a large-scale, free-surface water channel with the presence and absence of an endplate to represent a symmetry plane and to prevent free-surface effects. To obtain the same aspect ratios without using the endplate, appropriately sized wings were also used in a fully submerged arrangement. The velocity fields were acquired in the near wake of the flat plate using a two-dimensional (2D) two-component digital particle image velocimetry system, and simultaneous force measurements were performed. The wing had an aspect ratio ranging from 2 to 8 and pitches up to 45° in 1 or 6 convective times. The force measurement results revealed that major difference occurred after the motion was completed in which the lift and drag forces exhibited a local increase and decrease. Interactions between leading edge, trailing edge, and tip vortices were investigated; flow structures were correlated with the force-time histories, and variations with aspect ratio were also addressed.

A rapid path planning (RPP) method that fast generates suboptimal zero propellant maneuver (ZPM) paths was proposed previously, and to further enhance the planning performance, the differential flatness of the model is studied in this paper. For the crucial issue of singularity in flatness planning, the equivalence between the state transformation singularity and the control transformation singularity is proved for the feedback-linearizable multi-input multi-output (MIMO) affine system, which simplifies the singularity investigation of the ZPM problem. The question of searching suitable flat outputs for specific types of large-angle maneuvers is raised before path planning. Four sets of flat output candidates, which are biased to different types of maneuvers, are examined, and it is shown that only the large-angle yaw maneuver is achievable. The essentially singular attitudes (ESAs), for which no outputs could linearize the model, are presented, and it is found that the large-angle pitch or roll maneuver path planning easily encounters singularity in the flat output space no matter what output functions are constructed. To the regular mission of large-angle yaw maneuvers, the flatness is utilized in the RPP method. A simultaneous reconstruction method is proposed to solve the large amount of implicit equations in the planning, and rapid planning for the ZPM path is achieved.

Accurately predicting power consumption of the coring bit is a crucial building block of a lunar sampling mission. The total power consumed by a coring bit consists of two portions: the power consumed by the cutters in drilling soil and the power required to overcome the friction between the soil and the bit body. Based on soil mechanics and the elastic–plastic theory, this paper proposes a model that describes the interaction between the cutter and the soil. Analysis of this model confirms the hypothesis that cutting speed has a minimal effect on the cutting force, especially when cutting speed is low. Based on the assumption that low-speed rotary cutting is similar to linear cutting, cutting force and power consumption are tested experimentally. Results show that power consumed by drilling soil contributes to 70% of the total power consumption, and the power consumed to overcome friction contributes to 30% of the total power consumption.

In the present study, two new shear deformation theories, trigonometric deformation theory (TDT) and trigonometric-hyperbolic deformation theory (THDT), are proposed. These theories are examined in detail for their effectiveness in static response of laminated and three-dimensional (3D) braided composite plates. Both models are based upon a shear strain shape function that yields nonlinear distribution of transverse shear stresses and also satisfies the traction-free boundary conditions on top and bottom surfaces of the plate. Thus, requirement of shear correction factor vanishes. Virtual work principle is used to obtain the governing differential equations and boundary conditions. These models are formulated and validated for the static responses of laminated and braided composite plates. A variety of numerical examples are analyzed. Comparison of various higher order theories with the proposed theories are performed. It is observed that both proposed theories are efficient and accurate for the static analysis of laminated and 3D braided composite plates. Intensive numerical studies of 3D braided composite are performed in detail. It is further observed that the geometric parameter (aspect ratio), boundary condition, fiber volume fraction, and braiding angle have a significant effect on the static response of the braided composite plates. In the framework of finite element analysis, both proposed theories anticipate exemplary results for the laminated and braided composite plates compared to existing theories.

This paper proposes a computationally efficient attitude control strategy for nanospacecraft fast reorientation maneuvers. The paper considers a 3U CubeSat for visual Earth observation missions with deployable solar panels, equipped with three reaction wheels, three magnetorquers, and a miniature star imager, due to a 0.1° stringent targeting requirement of the payload. The star imager is very accurate, but operational only at very small angular rates. Hence it cannot be used for attitude measurement during fast slewing maneuvers. The proposed attitude control strategy overcomes this limitation by implementing a combination of open-loop and closed-loop control schemes based on the simplifying assumption of negligible gyroscopic torques. This leads to a straightforward onboard computation of the control actions required by the reaction wheels, within the limits imposed by saturation, without solving complex and computationally intensive time-optimal solutions onboard, which would not be compatible with CubeSats. The open-loop control phase is followed by an accurate closed-loop phase, accurately pointing the spacecraft toward the target orientation. The strategy is validated by numerical simulations, including robustness with respect to system uncertainties, by Monte Carlo analysis.

A system for supervision of aircraft status was developed, modeled and simulated. The system assesses overall aircraft status based on data available from external (out-of-aircraft) and internal (aircraft systems and crew) sources. The paper describes system requirements, functions, and architecture. It includes an example of an application. The aircraft and its systems status are assessed using the analytic hierarchy process (AHP) method, structured in a multicriteria architecture, which takes into account both quantitative and qualitative factors. To illustrate the proposed approach, a sample aircraft supervision system is developed to prove AHP applicability and to illustrate how this method could be implemented. To check aircraft sensor data, the proposed methodology uses the aircraft flight mechanic model. The developed supervision system shows the aircraft/systems status in a clear, quantitative way.

Bearing vibration signal separation is essential for fault detection of gearboxes, especially where the vibration is nonstationary, susceptible to background noise, and subjected to an arduous transmission path from the source to the receiver. This paper presents a methodology for improving fault detection via a series of vibration signal processing techniques, including signal separation, synchronous averaging (SA), spectral kurtosis (SK), and envelope analysis. These techniques have been tested on experimentally obtained vibration data acquired from the transmission system of a CS-29 Category A helicopter gearbox operating under different bearing damage conditions. Results showed successful enhancement of bearing fault detection on the second planetary stage of the gearbox

The present research proposed a method to reduce the Gurney-flap drag penalty with a dielectric barrier discharge (DBD) plasma actuator, which was mounted in front of the Gurney flap to generate a countercurrent wall jet. The research was based on numerical simulation, which was performed on a NACA 0012 airfoil with Gurney-flap heights (h) of 1, 2, and 3%c at ${\mathsf{R}}_c=6.84\times {10}^5$. According to the simulation results, the wall jet was able to decrease the additional drag on the Gurney flap by lowering the upstream pressure of the Gurney flap without disturbing the flow field downstream of the Gurney flap. The drag reduction effect was more obvious at low angles of attack and small Gurney-flap height, and a drag reduction of 15% was achieved at $\alpha =0°$ with 1%c flap height. This method also lowered total lift, but it was not strong enough to impair the high-lift property of the Gurney flap. For all tested Gurney-flap heights, the lift-drag ratios under low to moderate $C_L$ were enhanced by this approach. Comparing with the clean airfoil, a 12% increment of lift-drag ratio was obtained at cruise with plasma control at $h/c=1\%$ flap height.

This research is aimed at developing an efficient online path planner for unmanned air vehicle guidance in completely unknown three-dimensional (3D) rough terrain environments. A novel algorithm is proposed that directly incorporates the vehicle dynamics in the guidance strategy. A suitable point mass dynamic model is also developed. The flight path forms gradually as a result of applying the guidance commands to the vehicle dynamics. A key feature of this approach is real-time assessment of terrain characteristics and using this information in the guidance procedure. The problem is considered within a fuzzy behavior–based framework. The guidance algorithm uses acquired information from the onboard sensors and rapidly issues commands that will guide the vehicle safely to an intermediate position within the sensor range. Two behaviors are introduced: go to target and 3D terrain following/terrain avoidance. The issued commands are then integrated with adjustable weighting factors. Simulation results demonstrate a significant enhancement in vehicle autonomy level. Intelligent decision-making capability afforded by this approach allows for autonomous and safe low-level flight in mountainous areas.

A smart material–based active external trailing-edge flap was designed for wind turbine blades in this study. Its vibration-fatigue load reduction potential was evaluated during normal operation. The blade–external flap configuration was chosen based on a wind tunnel–testing report of the National Advisory Committee for Aeronautics. The internal structure of the flap was determined based on an aeroelastic sectional analysis code, with the input of aerodynamic forces obtained from a computational fluid dynamics simulation in order to meet strength and serviceability requirements. The work done by the aerodynamic forces and centrifugal forces acting on the flap was calculated to determine the required actuator capacity. Piezoelectric stack actuators combined with a double-lever mechanical linkage were designed to achieve the required force and displacement in order to activate the flap. By combining the flap control scheme and the aerodynamic-aeroelastic simulation, the time responses of the wind turbine blades were obtained. The fatigue analysis showed that the external flap can reduce the fluctuation of the blade root flapwise bending moment and the magnitude of the damage equivalent loads.

This paper reports an experimental study on the combined effect of drain port eccentricity and port diameter on air core vortex formation in a cylindrical tank with a rotating liquid column. The study is motivated by the need to control air core vortex formation in applications like propellant feed systems of rocket engines. Based on a series of experiments in a water tank with varying discharge port diameter and port eccentricity values, the study shows how the vortex formation can be suppressed using an appropriate combination of diameter and eccentricity of the discharge port. Experiments are carried out over a range of values of initial speed of rotation (60–200 rpm) of the water column. The height of the water column at which the air core vortex extends to the discharge port, is used as a measure of the vortex growth rate. The study proposes a correlation between the critical port diameter (which suppresses vortex formation) and port eccentricity. The intermittency of vortex formation in the eccentric port configuration is also studied using flow visualization.

In this paper, flame stabilization and propagation inside a kerosene-fueled two-stage strut dual-mode scramjet combustor is experimentally investigated at Mach 2 and Mach 3 with the stagnation temperature of 958, 1,163, and 1,626°C (1,231, 1,436, and 1,899 K). A $\mathrm{2,000}\mathrm{f}\mathrm{r}\mathrm{a}\mathrm{m}\mathrm{e}\mathrm{s}/\mathrm{s}$ high-speed imaging camera was used to record the flame structure during the combustion. The results show that when the scramjet is operating at lower stagnation temperatures of 958 and 1,163°C (1,231 and 1,436 K), the flame flashback equivalence ratio is much higher than the flame blowout limit, and the upstream strut can get a stabilized flame once it is reignited. A higher stagnation temperature and a lower inflow Mach number are advantageous to trigger flame flashback. When the scramjet is operating at a higher stagnation temperature of 1,626°C (1,899 K), continuous and oscillatory change of the flame and the axial wall pressure profiles were obtained.

The objective of this work is to reveal the significance of a trailing-edge attachment on the flow structure over a generic nonslender delta wing using the dye visualization technique on the top-view plane. Instantaneous images are acquired by stereoscopic particle image velocimetry (stereo-PIV) to calculate time-averaged flow data, with angle of attack, $\alpha$, and yaw angle, $\theta$, are varied within $7°\le \alpha \le 17°$ and $0°\le \theta \le 15°$, respectively. It is shown that time-mean locations of vortex breakdown for the nonslender delta and lambda wings occur in the field close to the wings. As a result, no remarkable alteration is observed for the spatial locations of the vortex breakdown. The experiments also show that the $\theta$, which is a crucial parameter, alters the flow structure over the nonslender delta wing substantially compared with the lambda wing. It is concluded that the trailing-edge attachment plays vital role when $\theta$ becomes effective.

In this paper, the authors describe a software they have developed for wing–body parametric modeling and high-lift system design. This software has the following features: (1) it generates and modifies geometric as well as aerodynamic analysis in the conceptual and preliminary aircraft design phases; (2) it integrates high-fidelity computer-aided design (CAD) and computational fluid dynamics (CFD) software by a graphical user interface(GUI); and (3) it has an aerodynamic/mechanical integrated design and kinematic simulation for a three-dimensional (3D) high-lift device. Based on the application program interface (API) techniques of specialized third-party software, the software integrates B-spline fitting and modeling of airfoil and complex curves, modeling of the fuselage and wings, a high-lift system design, the automatic generation of a structured grid, and high-fidelity CFD code based on Reynolds-averaged Navier–Stokes (RANS) equations. First, the geometry of the fuselage and wing is parametrically generated by lofting with control curves, and then the initial graphics exchange specification (IGES) format geometry is exported to an automatic grid generation module. Finally, the structured grid file is exported to the CFD solver for aerodynamic analysis. The NACA0012 airfoil and DLR-F6 FX2B wing–body model that was published at a workshop are selected for software validation. The cruise configuration model (with nacelle, wingtip, and empennage designed manually) of a 150 seat airliner designed using this software was subsequently used for a high-lift aerodynamic/mechanical integrated design and kinematic simulation. The results illustrate that the software was able to integrate wing–body modeling and high-lift aerodynamic/mechanism design in the conceptual and preliminary aircraft design phases.