Journal of Aerospace Engineering

A nonlinear dynamic modeling method is developed for a deployment and locking mechanism composed of laminated composite appendage. Unlike most formulations of linear models which ignore coupled and nonlinear terms resulting in a seriously improper response, the present model takes into account the effects of geometric nonlinearity and coupled deformations. In order to accurately obtain the dynamic response of laminated composited appendages, nonlinear strain-displacement relations for laminated plate/shell elements are presented, and the corresponding formulations are derived from the Piola–Kirchhoff stress tensor for evaluating the internal forces. Furthermore, the effect of contact and impact located at a spring hinge is investigated, which can achieve the actuating and locking functions. To study the dynamic behavior of contact-impact, the generalized contact-impact forces between the pin and locking groove are considered in the model. Meanwhile, the Jacobian matrices of geometric and momentum constraints are derived from hinge kinematics. Finally, the complete expressions including coupled deformation terms, nonlinear stiffness, additional stiffness terms, and contact-impact forces are presented, and a full analysis is achieved by using both the linear model and nonlinear model, respectively. Numerical simulation results are obtained to verify the dynamic effects of coupled deformation terms, the nonlinear stiffness and additional stiffness terms in the present model.

The main emphasis of the present study is to discuss the effect of chemical reaction on flow, heat, and mass transfer of Williamson nanofluid over a stretching sheet. The flow is considered under the influence of nonlinear thermal radiation. To obtain the numerical results, the governing equations have been reduced to a set of nonlinear ordinary differential equations using appropriate similarity transformations. An efficient Runge-Kutta-Felhberg 45 order method along with a shooting technique is used to get the required solutions. The accuracy of the method used is verified with the existing results, and they are found to be in good agreement. Obtained numerical solutions are presented in the form of graphs and tables for a various range of slip boundary condition and for different values of flow pertinent parameters, such as permeability parameter, radiation parameter, Lewis number, Heat capacities ratio, diffusivity ratio, Schmidt number, chemical reaction parameter, volumetric volume expansion coefficient, and Prandtl number. Finally, the outcome of the problem is written in the form of a conclusion based on the tables and plotted graphs.

Fast nonsingular terminal sliding mode controllers (FNTSMCs) are proposed for underactuated spacecraft formation reconfiguration without either the radial or the in-track thrust. A nonlinear dynamic model of underactuated spacecraft formation is developed, which is then linearized about circular reference orbits. Based on the linearized model, the system controllability is analyzed for either underactuated case. Due to the loss of control in a certain direction, the disturbances consisting of the linearization errors and external perturbations do not enter the system from the same channels as those of the control inputs, and therefore turn into unmatched disturbances. To achieve underactuated formation reconfiguration in the presence of unmatched disturbances, underactuated FNTSMCs are designed to indirectly render the system states convergent to the neighborhood of equilibrium by using the inherent coupling of these system states. Modified FNTSMCs are also presented to enhance the system performance in case of a sharp increase in the control inputs when the system states are away from the sliding surfaces. Meanwhile, the terminal convergent region of each system state is presented. The overall stability of the closed-loop system is guaranteed by a Lyapunov-based approach. Theoretical analyses are validated by numerical simulations that formation reconfiguration could still be achieved by the proposed controllers even with the loss of radial or in-track control.

The diversion effect of a conic flame deflector for a lunar module ascent stage is investigated experimentally and numerically. Eighteen tests were performed in a vacuum chamber mounted with a liquid helium cryogenic pump using a scaled engine and models. The experimental results indicate that the conic flame deflector can guide the exhaust gas away from the ascent stage model more effectively than a plate; meanwhile, the conic flame deflector can create a reverse torque, which helps to correct the deflection of the ascent stage model. Finally, the experimental results show that the reverse torque changes with the distance and angle between the ascent stage model and the conic flame deflector model. Numerical simulations have been done with the computational fluid dynamics and direct simulation Monte Carlo (CFD/DSMC) method, and the numerical results are in good agreement with the experimental results. The numerical results reveal that the reason for the above phenomena is shock waves. An expansion wave starting from the nozzle lip and a compression wave starting from the cone surface control the amount and speed of the exhaust gas by changing the magnitude and direction of flow velocity and then influence the aerodynamic pressure on the ascent stage model.

Earthbound testing of the mobility of lunar, Martian, and other extraterrestrial rovers benefits from the use of suitable soil simulants. To this end, a granular material called Fillite was selected as a simulant for modeling high-sinkage, high-slip situations that could be encountered by rovers, such as the one encountered by the Spirit rover on Mars. Fillite consists of alumino-silicate hollow microspheres harvested from the pulverized fuel ash of coal-fired power plants. It is available in large quantities at a reasonable cost and it is chemically inert. The focus of this paper is to summarize geotechnical characterization of Fillite, specifically the mechanical properties such as shear strength parameters, elastic modulus, Poisson’s ratio, and small-strain shear modulus. These measured properties are expected to enable analysis of rover mobility tests conducted in Fillite. The properties of Fillite are compared with the known and estimated properties of Martian and lunar regoliths as well as of other commonly used simulants. Fillite is quite dilatant. The peak and critical angles of internal friction of Fillite are smaller than those of most other simulants. Smaller shear strength, coupled with much smaller bulk unit weight as compared to other simulants, would result in smaller bearing and shearing resistances. This is expected to allow for better simulation of the intended high-sinkage, high-slip environment for rover mobility studies.

Tensile integrity (tensegrity, tensegric) structures are suggested for possible use in architectural works, bridges, covering large areas, and especially for habitats in outer space and on space bodies because of their minimal need for structural materials and ease of transfer and mounting. Structurally, their behavior is nonlinear and thus their analysis necessitates special care with recourse to special techniques. This study shows that their structural analysis can successfully be performed by a method called total potential optimization using metaheuristic algorithms (TPO/MA) with no special precautions in a very general way. In the applications, which are conducted using software prepared for analyzing all types of nonlinear trusses and trusslike structures, the load-displacement behaviors of tensegric structures are investigated to include also the effect of prestressing level on cables. Genetic algorithms are used in this study as the metaheuristic technique for solving the optimization part of the problem.

A mathematical study is presented to investigate the influence of variable transport properties, momentum, thermal, and mass slip on magnetohydrodynamic (MHD) momentum, heat, and mass transfer in a Darcian porous medium. Slip effects are simulated via careful imposition of boundary conditions at the wall. Joule heating and viscous dissipation are also studied. The governing partial differential boundary layer equations are analyzed using Lie group theory and rendered with appropriate transformations into a system of nonlinear, coupled ordinary differential equations. The multiphysical boundary value problem is dictated by 12 thermophysical parameters: concentration diffusivity ($D_c$), Hartmann magnetic number ($M$), permeability ($\mathrm{\Omega }$), Eckert number (Ec), momentum slip ($a$), thermal slip ($b$), mass (species) slip ($d$), Prandtl number ($\mathsf{Pr}$), Schmidt number ($\mathsf{Sc}$), power law index ($m$) for nonisothermal and nonisosolutal effects, viscosity variation ($A$), and thermal conductivity variation ($S$). A numerical solution is obtained for the effects of selected parameters on transport characteristics using the robust Runge-Kutta-Fehlberg fourth-fifth order numerical method. Excellent correlation is achieved between the present computational results and for the constant transport properties ($A=S=Dc=0$), nonporous ($\mathrm{\Omega }=0$), nonthermal slip ($b=0$), nonsolutal slip ($d=0$), and nondissipative solutions without joule heating ($\mathrm{Ec}=0$). Increasing momentum slip enhances temperatures, whereas increasing thermal slip reduces them. An increase in thermal conductivity boosts temperatures, whereas greater viscosity reduces temperatures. Increasing the magnetic parameter suppresses velocity and increasing permeability parameter elevates temperatures. Species concentration is enhanced with increasing concentration diffusivity and permeability parameter but depressed with increasing viscosity. Furthermore, concentration is enhanced with momentum slip but reduced with mass slip parameter. Increasing magnetic field is observed to aid species diffusion in the regime. The present study finds applications in trickle bed reactor hydromagnetics, magnetic polymeric materials processing, and MHD energy generator slip flows.

The steady stagnation point flow of a micropolar fluid over a stretching/shrinking sheet with second-order velocity slip is studied. Similarity equations are obtained using similarity transformation, which are then solved numerically using MATLAB routine boundary value problem solver (bvp4c) based on the finite-difference method. Numerical results show that dual solutions exist for a certain range of the shrinking parameter. The dual solutions for velocity and microrotation distribution with first-order, second-order velocity slip parameter and micropolar parameter are shown graphically. It is observed that the range of the stretching/shrinking parameter for which the solution exists increases with the increase of the first-order slip parameter and micropolar parameter, whereas it decreases with the increase of the second-order slip parameter. The linear stability analysis of the obtained results was performed to show that the first solution branch is linearly stable, whereas the other is always unstable.

This study considers the attitude control problems of a quadrotor helicopter with actuator failures and external disturbance and thus proposes a self-repairing control scheme via adaptive control and quantum logic. An improved model reference adaptive compensation controller is designed to compensate for the failures and disturbance, as well as to guarantee that the attitude angles of the quadrotor helicopter track the state output of a given reference system in which the upper bounds of disturbance and the unparameterized stuck faults are unknown. The asymptotic state tracking properties of the designed adaptive system are proven using Lyapunov stability theory. In addition, quantum logic is introduced to increase the control accuracy of this method. Finally, a numerical simulation example is presented to demonstrate the validity of the proposed control scheme.

Meeting future goals for aircraft and air traffic system performance will require a fundamental shift in approach to aircraft and engine design. In 2005, the National Aeronautics and Space Administration (NASA) released plans of a next generation commercial airplane for 2030 combining the blended wing body (BWB) and a superconducting distributed propulsion system. The BWB concept adapts NASA’s cruise-efficient short take-off and landing (CESTOL) airframe. The propulsion system employs distributed electric fans, which are embedded on the upper surface of the airframe, driven by superconducting motors with power provided by two wing-tip mounted turboelectric generators. This paper describes a method to design a turboelectric distributed propulsion (TeDP) system on the hybrid wing body airframe, including a way to obtain the propulsor number and its weight, a method to simulate boundary layer ingestion, and a method to calculate electric system performances and its weight. An examination of the system thermodynamic performance for a range of fan pressure ratio (FPR) was also made. A comparison with results from the NASA $N+3$ technical reports on the TeDP system has been performed. The new model returned similar results. Finally, the effects of intake distortion were evaluated. It has been found that the number of propulsors is impacted by the motor size, the span-wise length, and the inlet condition. The weight of the propulsors unit reduces by increasing the propulsor number. What is more, the ingesting boundary layer reduces the fuel consumption, and the benefits are sensitive to the inlet pressure loss.

In this paper, a modified laminated finite-element model based on subcell of unit cell and CDM material model was developed in nonlinear explicit finite-element code to investigate the dynamic response of 2D triaxial braided composite under ballistic impact of projectile. Six subcells, which were considered as laminate with different stacking sequences using solid elements, were employed to represent one unit cell of triaxial braided composites, therefore, the fiber-laying direction of braiding architecture is reproduced. In the prediction of the subcell model, an approximate round conical deformation area forms in the target and the failure modes are mainly longitudinal and transverse cracking due to tension. Compared with the continuum model, the subcell model is capable of analyzing the energy absorption of fiber in each direction. Although the failure modes are similar for both methods, the subcell model is more effective in prediction of impact wave propagation than the continuum model because it describes the fiber arrangement of triaxial braided composites. The developed subcell model will be an effective tool for the design process of carbon/epoxy composite fan casings.

In this paper, the interface debonding and slipping of carbon fiber-reinforced ceramic-matrix composites (CMCs) subjected to different fatigue loading sequences have been investigated using the micromechanics approach. There are two different types of fatigue loading sequences considered: (1) cyclic loading under low peak stress for N1 cycles, and then high peak stress for N2 cycles; and (2) cyclic loading under high peak stress for N1 cycles, and then low peak stress for N2 cycles. Based on the fatigue damage mechanism of fiber slipping relative to matrix upon unloading/reloading, the interface debonded and slip lengths are determined by fracture mechanics approach. The relationships between interface debonding, interface slipping, interface wear, cycle number, fatigue peak stress, and fatigue loading sequence have been determined. The effects of peak stress level, interface wear, cycle number, and loading sequence on the interface debonding and slipping of fiber-reinforced CMCs have been analyzed. With increasing cycle number and the peak stress level, the interface debonding and slipping range increase, leading to the increase of unloading residual strain and hysteresis loops area, and the decrease of hysteresis loops modulus of fiber-reinforced CMCs. The cyclic fatigue hysteresis loops of unidirectional C/SiC composite under multiple fatigue peak stress levels of ${\sigma }_{\max }=200$, 220, and 240 MPa have been predicted.

A finite volume–based unstructured inviscid flow solver has been developed for the study of nonequilibrium effects in high enthalpy flows. Four different inviscid flux computation schemes are tested for literature reported test cases like unsteady wave motion in a shock tube and high enthalpy flow over cylinder/sphere. During the present assessment, encouraging agreement has been noticed in capturing all the essential flow features in the reacting media. Here, the Rusanov scheme is noticed to be computationally cheaper among all the schemes with noticeable diffusion around strong gradients. Shock tube test case revealed the existence of various shock Mach numbers for a given driving pressure ratio in the presence of reactions. Similarly, freestream Mach number is found to have limitations in representing the shock standoff distance in case of high enthalpy flow over cylinder and sphere. In line with this, real gas effects are prominently observed for the case of the cylinder in comparison with the sphere while predicting the complete shock shape using Billig’s correlation. In all, present investigations portray the limitations of conventional nondimensional numbers in predicting the flow features and flow properties in the presence of reactions in all the dimensionalities.

A number of numerical predictions are performed to evaluate the aerodynamic and infrared radiation characteristics of spherical convergent flap nozzles (SCFN). The accuracy of the computations was verified by comparing the calculated wall pressure values to experimentally measured ones. The prediction results show that the vortices produced by the secondary flow at the corner of the divergent section are the main factor to produce thrust loss without any vectoring action. Compared to the axisymmetric nozzle, the SCFN caused up to 1.86% thrust loss and 3% flow rate loss. The yaw angle had no effect on the inner flow field in the divergent section and the exhaust flow because the yaw vector turning is achieved by gimballed subsonic flow. The maximum thrust losses were only 1 and 0.69%, which were caused by the pitch and yaw vector, respectively. The aerodynamic vector angle was approximately consistent with the turning angle due to geometric deflection of SCFN. With a 15° geometric pitch angle, the maximum radiation was depressed by 8.76% on the vertical plane and by 23.9% on the horizontal plane relative to that without pitch. With a 10° geometric yaw angle, the maximum radiation was depressed by 11.8 and 5.7 on the horizontal and vertical planes, respectively.

This paper presents an adaptive self-tuning flight control system (AST-FCS) for a fixed-wing aircraft, applied to a high-fidelity dynamic model of the ATLANTE unmanned aircraft system (UAS). This new adaptive control system is based on adaptive predictive (AP) control methodology and a new guidance system. AP control has been implemented in many complex industrial applications with excellent results, but it has not been applied in the aerospace sector yet. Experimental results obtained in simulation by flying a high-fidelity model of the ATLANTE UAS illustrate excellent control system performance and show the robustness of the AST-FCS, which has been designed without using accurate knowledge of the aircraft dynamics embedded within the model. It is expected that the application of AP control in this field could introduce significant advantages in aircraft control, improving performance in a wide variety of flight conditions and reducing design costs.

Drilling and coring, due to their efficient penetrating and cutting removal characteristics, have been widely applied to planetary sampling and returning missions. In most autonomous planetary drilling, there are not enough prior seismic surveys on sampling sites’ geological information. Sampling drills may encounter uncertain formations of significant differences in mechanical properties. Additionally, given limited orbital resources, sampling drills may have a stuck fault under inappropriate drilling parameters. Hence, it is necessary to develop a real-time drilling strategy that can recognize current drilling conditions effectively and switch to appropriate drilling parameters correspondingly. A concept of planetary regolith drillability based on the rate of penetration (RoP) is proposed to evaluate the difficulty of the drilling process. By classifying different drilling media into several drillability levels, the difficulty level of drilling conditions can be easily acquired. A pattern recognition method of support vector machines (SVMs) is adopted to recognize drillability levels. Next, a set of suitable drilling parameters is tuned online to match the recognized drilling conditions. A multilayered simulant drilling test indicates that this drilling strategy based on drillability recognition can identify different drilling conditions accurately and have good environmental adaptability.

A numerical study is conducted to investigate the effects of the inflow Knudsen number ($Kn$) on hypersonic chemical nonequilibrium rarefied flows using the unstructured direct simulation Monte-Carlo (DSMC) method, and an empirical formula for fast prediction of flow-field structure for different values of $Kn$ is proposed and verified. First, the flow around a vertical plate is analyzed at a free-stream velocity of $\mathrm{7,500}\mathrm{m}/\mathrm{s}$ using a perfect-gas model with inflow $Kn$ values ranging from 0.035 to 13.36. Through analysis of the flow-field characteristics with varying $Kn$ values, a linear expansion law of flow-field versus the square root of $Kn$ is discovered and an empirical formula based on least-squares fitting is developed for fast prediction of the flow structure at different values of $Kn$. Further, the effects of chemical nonequilibrium on hypersonic rarefied flow-field structure are investigated to verify the applicability of the empirical formula. Finally, cases of the cylinder flow are simulated in the regimes from continuum flow to free-molecular flow using a five-species chemical reaction model. The flow-field characteristics with varying $Kn$ values, with the chemical nonequilibrium model, agrees well with those for perfect-gas conditions. It is shown that the empirical formula is also suitable for the use of flow with chemical reactions. The shock position in a continuum-flow case predicted by the empirical method agrees well with data from the literature. The applicability and accuracy of this method over a wide range of $Kn$ have been verified.

The existence of a liquid water ocean under the icy surface of Jupiter’s moon Europa has been considered highly probable for decades. It has been speculated that such an environment could host extraterrestrial life. However, the thickness of Europa’s top layer of ice and the depth of the ocean underneath are currently unknown. The authors investigate the feasibility of using seismic-inversion analysis to estimate the thickness of the ice shell and the depth of the subsurface ocean in a hypothetical event, in which a spacecraft-released impactor makes a planned collision with Europa’s surface. The wave motions on Europa’s surface that would result from a planned impact could be measured by impactor-probes penetrated into the ice shell. This work uses finite element method (FEM) modeling of seismic wave propagation to compute wave motions with a guessed ice thickness and ocean depth as input. Inverse modeling is used to determine the input that produces the best results when comparing the computed wave motions attributable to the input to the wave motion owing to pseudotarget values for ice thickness and ocean depth. To arrive at the solution of the inverse problem, the genetic algorithm (GA) was employed. The presented numerical simulation demonstrates the potential of using the inverse-modeling method for estimating the thickness of Europa’s ice shell.

The present study attempts to fix the location of thermal throat by employing transverse fuel injection downstream and adding a short constant-area section in the divergent part in a realistic scramjet combustor configuration in order to depress the pressure rise effects on the isolator. Ground tests using a direct-connect model along with pressure measurements and three-dimensional numerical simulation are also performed for the present investigations. The results indicate that the thermal throat is fixed in the constant-area section for the configuration with a short constant-area section in the divergent part. Numerical simulations also indicate that the constant-area section is helpful for ignition and flame holding. The configuration with a short constant-area section increase heat release in the combustor and decrease entropy production. Numerical simulation results imply that there could exist more than one thermal throat in dual-mode scramjet combustor, and the constant-area section in a divergent part could reduce the number of thermal throat.

The development of upscaled wind turbines of tomorrow would benefit from a rationale change in load-control strategies. Active flow-control devices attached to existing blade designs constitute an interesting alternative. This study presents a numerical assessment of the aerodynamic characteristics of airfoil sections specifically intended for wind turbine applications, with modular flow-control devices attached. Based on the slotted trailing-edge flap concept, these modular devices could be appended to an existing blade design, providing a cost-effective method of active control with low energy of actuation. Their use also requires minimal design modifications to the original blade and retooling of its manufacturing process. Numerical results are presented for the study of the effects of slotted flaps applied to two airfoil sections, the NACA ${64}_3-618$ and the DU 93-W-210, which are currently used on the NREL-5MW reference wind turbine. By adding new knowledge about the effects of flaps applied to airfoil sections specifically intended for wind turbine applications, this study can provide useful data for a wider research community in developing new designs and load mitigation strategies for wind turbine rotors.

The objective of this paper is to numerically investigate the three-dimensional side loads and associated flow physics of a conical nozzle during the fire-in-the-hole staging event of the multistage rocket. The transient flow in nozzle is examined by a three-dimensional time-accurate numerical method with the start-up process of the upper-stage motor and the prescribed nominal separating motion of the lower stage. The impacts of asymmetric lateral motion due to non-nominal staging movement on the nozzle side loads are also numerically analyzed. These motion deviations include the relative angle between axes of two stages, relative lateral displacement between two stages, superposition of relative angle with relative lateral displacement, and angle of thrust vector control. As presented in the simulation results, the flow separation in nozzle is significantly prolonged and intensified by the obstruction of the lower stage. With the location of shockwave reflection approaching the dome of lower stage, the nozzle flow structure experiences two different, stable free shock separation (FSS) patterns. The transition from the first pattern to the second pattern results in the flow oscillation and the side loads peak. The staging motion deviations result in not only the enlarged nozzle side loads, but also a sustained unidirectional effect that may influence the flight stability of the upper stage. An essential ingredient driving the flow unsteadiness and asymmetry is related to the asymmetric reverse supersonic flow deflected by the dome and skirt corner of lower stage. Lastly, the flight data of maximum lateral force during staging of Scout’s stage I and stage II is discussed and compared with the computed results in the article.

To provide a basic model for future lunar explorations, the compressional (P) and shear (S) wave velocity profiles in the shallow lunar subsurface are deduced from the velocities in a lunar soil simulant called FJS-1. The P-wave and S-wave velocities in several types of samples derived from FJS-1 were measured by laboratory bender element tests to examine the effect of confining stress, void ratio, and fine fraction content. Because of its high fine fraction content, FJS-1 exhibited remarkably lower velocities than other materials, although they were similar in terms of dependence of the confining stress and void ratio. Based on the test results, empirical equations for deriving the velocities of FJS-1 were proposed as functions of confining stress and void ratio. The equations were then used to deduce the P-wave and S-wave velocities in the shallow lunar subsurface. Two sets of continuous velocity profiles up to a depth of 60 m were estimated and found to be in agreement with existing velocity data derived from both measurement and estimation.

Poor lateral-directional flight quality is a critical technical problem in flying wing aircraft design. This paper proposes an approach to achieve lateral-directional dynamic stability of flying wing aircraft. This approach is different from general methods such as installing a wingtip vertical stabilizer or depending on a stability augmentation system. Through directly including the judgment of aircraft dynamic stability into the aerodynamic configuration design, sufficient lateral-directional stability can be guaranteed without relying on other systems. A lateral-directional dynamic stability satisfaction judgment function was built based on MIL-F-8785C. The lateral-directional dynamic stability was improved using only the adjustment of the spanwise dihedral layout. To validate the feasibility of this design method, this study selected a flying wing aircraft with a high aspect ratio as the object. Two optimized configurations with different aerodynamic shape but similar stability characters were finally obtained. Each configuration achieved flight quality which surpassed Level 2 requirements of MIL-F-8785C in the entire airspeed range.

Most spacecraft have at least one pressurized vessel on board. For robotic spacecraft, it is usually a liquid propellant tank or battery. For human spacecraft, there are also pressurized living quarters and life-support systems. One of the design considerations of such spacecraft is the possible damage that might occur in the event of an on-orbit impact by a micrometeoroid or orbital debris (MMOD) particle. While considerable energy and effort has been expended in the study of the response of nonpressurized spacecraft components to these kinds of impacts, relatively few studies have been conducted on the pressurized elements of such spacecraft. In addition, the design criteria currently used by the National Aeronautics and Space Administration (NASA) for pressurized tanks operating in the MMOD environment have not been tested or scrutinized since they were first proposed nearly 45 years ago. This paper reviews current NASA design criteria for pressurized vessels and offers suggestions for next steps in their further development.