Journal of Aerospace Engineering

Asymmetric dielectric barrier discharge plasma generating a wall-bounded jet without mechanical moving parts was studied as a flow control device. A reliable model providing thrust value of the plasma actuator is difficult to develop without incurring a heavy calculation cost attributable to its unsteady characteristic and asymmetric electric field distribution. In this paper, a new semiempirical thrust prediction model was developed based on the one-dimensional electrohydrodynamic effect. It is implemented in a computational fluid dynamics solver as a body force term. We determined that the thrust of the plasma actuator is proportional to the energy consumed in the actuator, which is regarded as a capacitor in the alternating current (AC) circuit. An analytic estimation of the capacitance of the actuator is performed, and the sensitive design parameters dielectric constant, thickness and the upper electrode thickness are considered in the capacitance value. The thrust value from the model is inserted as a body force term in a Navier-Stokes equation solver and the body-force region is specified by the extent of the region of the discharged plasma. Grid dependence of the model is verified and the velocity profile changed by the actuator is compared with previous reference experiment data to validate it. Thus, the authors conclude that the developed model can provide the proper thrust value and a two-dimensional velocity profile without incurring a heavy calculation cost.

In this paper, the authors present a procedure for the global solution by the use of evolutionary algorithms of trajectory optimization and parameter estimation problems for rotorcraft vehicles. The proposed approach makes use of a novel repair heuristic based on a sequential quadratic programming method to handle problem constraints. The resulting formulation effectively deals with the complex, highly constrained, optimization and optimal control problems plagued by multiple local minima or maxima that are often encountered in practical rotorcraft flight mechanics applications. The performance of the proposed procedures is assessed with applications dealing with the design of optimal inputs for the estimation of model parameters, with the determination of the optimal way of flying continued take-offs under category A requirements, and with the determination of the most dangerous areas of the height-velocity (H-V) plane in terms of impact velocity with the ground.

This paper presents a smart way to actuate and control the airfoil shape of a morphing wing in the open-loop architecture. The actuation system uses smart material actuators such as shape memory alloys, disposed in two parallel actuation lines, and its control is performed with a Mamdani-type fuzzy logic proportional derivative (PD) controller. The morphing wing project description, its actuation system structure, and the control design and validation are highlighted in this paper. The results obtained by both numerical simulation and experimental validation (bench tests and wind-tunnel tests) are presented as part of the control design and validation. An analysis of the wind flow characteristics, based on the information provided by the pressure sensors mounted on the flexible skin of the morphing wing in the wind-tunnel tests, is included as a supplementary validation; the transition between laminar and turbulent flows is real-time visualized, and the aerodynamic efficiency of the controlled morphing wing is evaluated.

In this study, a robust extended Kalman filter (EKF) for estimation of a spacecraft’s position and velocity by single station antenna tracking data against the effect of measurement matrix singularity is proposed. The designed filtering algorithm estimates the position and velocity of the spacecraft on direct nonlinear range-azimuth-elevation measurements. The proposed robust EKF is examined through the geostationary satellite orbital motion simulations and the recommendations on its application for various spacecraft missions are given.

Structural stiffness has been increasingly considered as a performance index or parameter indicating condition of structures in the fields of aerospace, civil, and mechanical engineering. In this paper, theoretical and experimental structural stiffness models are proposed, and random vibration responses of the structure are evaluated. As a demonstration, the proposed method is applied to a one-story frame structure generally used in many engineering applications, and the experimental results show that the structural stiffness increases as the thickness of girder augments. Due to the idealization and perfect column-girder joint conditions, the theoretically predicted structural stiffness value is always larger than that measured from the experiment. Since the experimental method includes all practical factors, the experimental values of structural stiffness represent real global stiffness of the structure. An empirical equation of structural stiffness for the one-story frame structure is thus obtained by fitting the experimental data. The experimental random vibration experiment is conducted and it demonstrates that the random vibration responses under the same excitation all decrease when the girder becomes thicker. The influence of crack damage in columns of the one-story frame on structural global stiffness and random vibration response is also experimentally investigated, and it shows that the structural stiffness reduces slightly when damage is present. For the example given where the damage locates at the midheight of columns and coincides with location of the inflection point, the local damage in the structure only imparts a little change on the structural global stiffness, and the random vibration responses of the intact and damage structures also exhibit little difference. The experimental structural stiffness model presented can be used to data-reduce the global structural stiffness from the random vibration experiment, and it can be in turn considered as a performance index to assess condition of or detect damage in the structures.

An open-source program implementing a block-structured adaptive mesh refinement method was adopted for the fine structure numerical simulation of detonation initiation in supersonic combustible mixtures. Simulations were conducted on a nested parallel computing system. The initiation process was specified as three stages, and their respective flow field characteristics were analyzed. Results indicate that a hot jet under specific conditions can have a similar effect as a pneumatic oblique bevel for inducing periodical shock-induced detonative combustion by a bow shock. The interaction of bow shock–induced combustion with the local detonation wave, produced by the reflection shock on the upper wall, can create a structure with two triple-wave points. The hot jet not only plays a role in the detonation initiation but also acts as a stabilizing control mechanism for detonation propagation. In the simulations in this study, the detonation wave propagates in an overdriven state initially and achieves self-sustaining motion after the shutdown of the hot jet. Subsequently, the final pisiform structure of typical stable Chapman-Jouguet detonation cells is formed.

In this paper, a new formation flying approach based on virtual structure is presented. In this architecture, formation control strategies are appropriate when a large number of unmanned aerial vehicles (UAVs) are involved. Control laws for formation control are designed based on both classical theory and inverse dynamics, and then a comparison study is performed. The effectiveness of the proposed control strategies is demonstrated through nonlinear 6-degrees of freedom (DOF) simulation.

This work investigates the aerodynamic interference among airplane components caused by rudder deflection for a typical turboprop aircraft geometry through the computational fluid dynamics technique. At no sideslip, an airplane is in symmetric flight conditions. The rudder deflection creates a local sideslip angle close to the vertical tailplane, and this effect is increased by fuselage and horizontal tail. Typical semiempirical methods, such as United States Air Force Stability and Control Data Compendium (USAF DATCOM), do not take into account for these effects, proposing the same corrective parameters both for pure sideslip and rudder deflection conditions. Numerical analyses executed on several aircraft configurations with different wing and horizontal tailplane positions show that the interference factors are smaller than those predicted by the USAF DATCOM procedure, providing guidelines for a more accurate aircraft directional control analysis and hence rudder preliminary design.

In this numerical study, the cooling of exhaust gas issued from a rocket engine in a ground test facility has been investigated by solving the three-dimensional governing equations. Simulations have been performed by employing discrete phase model with plain-orifice atomizer to cool the exhaust flow effectively by injecting the coolant (water) in the form of a fine spray. The effects of coolant flow rate, inlet gas temperature, injection pressure, injector diameter, gas cooler length, and distribution of water droplets on the cooling characteristics, such as the temperature at the gas cooler exit, droplet diameter, and percentage of unevaporated water have been discussed in detail. Simulations highlight that optimum particle diameter needs to be identified for effective cooling, and also uniform cooling of the exhaust gas is attained by employing staggered injector distribution (more injectors at the periphery compared to the core region). Predicted values of static temperature and pressure agree well with the experimental data obtained from a scaled-down model high-altitude test facility.

The cavity-strut supersonic combustion organization method has been found to improve the combustion performance of ethylene in recent studies. In this work, liquid kerosene combustion experiments have been carried out in a model scramjet combustor with a cavity and small strut upstream of the cavity. The effect of fuel injection allocation between the transverse injectors at the strut leading surfaces and duct wall upstream of the cavity on the combustion characteristics has been investigated. In the experiments, a larger bench thrust increment, a higher specific impulse, and an improvement in isolating the interference from the downstream region have been obtained as more of the fuel was injected from the strut injectors. Increasing the strut injection proportion makes the combustion zone translate to the divergent segment downstream of the cavity, and leads to the dominant flow changing from subsonic to supersonic in the minor fuel-rich condition.

The conventional concept for solar energy utilization in space is converting solar radiation to electricity and then using the electric power for various applications. For some applications, however, it is more efficient to utilize solar energy directly. Oxygen production from lunar regolith, for example, can be accomplished by using solar thermal power directly for thermochemical reduction of lunar oxides. Space-based plant growth can use photosynthetically active radiation (PAR) from solar spectra for biomass production and life support. The authors and their colleagues have been developing a new solar power system called the optical waveguide (OW) system for solar power utilization in space. In this system, solar radiation is collected by the concentrator array, which transfers the concentrated solar radiation to the OW transmission line made of low loss optical fibers. The OW transmission line directs the solar radiation to the site of solar power utilization. This paper discusses the technical background of the OW system and reviews development and testing of the engineering prototype of the OW solar thermal system during the National Aeronautics and Space Administration (NASA) in situ resource utilization (ISRU) analog test in 2010. Based on the results, performance and viability of the OW system in application to space solar power utilization are discussed.

An igloo-shaped magnesium structure, supported on a sintered regolith, covered by sandbags of regolith shielding, and made using rapid manufacturing technologies is proposed as a potential design of a lunar structure. Also, a thermal analysis is carried out to study the effect of the regolith shielding as well as the sensitivity of such designs to measurement uncertainties of regolith and sintered thermal properties. It is shown that 3-m-thick regolith shielding can insulate the structure regardless of its location on the Moon.

During the Apollo missions, seismic events with estimated body wave magnitudes of 5.5 were observed. Whereas the terrestrial equivalent can cause moderate to severe damage to buildings, the effects on a lunar structure is not the same, as the lunar geological structure is very different than that of the Earth. The many challenges in predicting the effects of an earthquake are known. Lunar seismic effects are a much greater challenge because of the severe lack of data. Therefore, the authors begin to consider lunar seismic effects to assess the vulnerability of designs that ignore seismicity. A seismic event of body wave magnitude 7 has been developed in this paper based on the best available data and applied to the typical structure. The results indicate that the risk associated with these events is low in part because such structures are designed with a relatively high factor of safety. This paper provides an extensive and specific technical discussion of lunar seismology as applicable to structural design.

A concurrent simulation and experimental validation for curvilinear fatigue crack growth (FCG) analysis under both constant amplitude and overload spectrums is proposed in this paper. The simulation methodology is based on a small time-scale fatigue crack growth model and the extended finite element method (XFEM) to calculate the stress intensity factor solution of an arbitrary curvilinear crack. Parametric studies are used to determine the algorithm parameters in the numerical fatigue crack growth simulation. Following this, experimental testing on modified compact specimens is performed under both constant amplitude and overload loadings for model validation. Experimentally measured crack growth orientations and lengths are compared with numerical simulations. Both the experimental and simulation results show the overload retardation behavior for curvilinear cracks under overload loadings. The investigated periodic overload loading has no significant impact on the crack growth orientations. Several conclusions and areas of future work are identified based on the proposed numerical and experimental investigations.

A velocity-free attitude fault tolerant control scheme is presented to stabilize spacecraft attitude. A terminal sliding-mode observer is developed to reconstruct system states. With the reconstructed attitude and angular velocity, a control law is synthesized without any knowledge on actuator faults. The controller guarantees all the signals of the closed-loop attitude system to be uniformly ultimately bounded even in the presence of actuator fault and external disturbance. The control power is rigorously ensured to be within the maximum magnitude of output torque. An illustrative example is presented to evaluate the control performance.

This paper discusses a simplified approach for estimation of skin temperature of aircraft at various flight conditions using computational fluid dynamics to aid precise estimation of cockpit heat load. Major heat load generated inside the cockpit is because of skin friction of air at high speeds. Hence, estimation of the stagnation temperature on aircraft skin is important. Accurate estimation of heat loads will enable the optimum design of air conditioning system with minimal weight and fuel penalty to improve endurance of aircraft. This paper gives a methodology for the calculation of skin temperature by carrying out three-dimensional (3D) modeling of the cockpit in computer aided design software, tetrahedral meshing and simulation in computational fluid dynamics (CFD) design software at various angles of attack (AoA), aircraft speed, and ambient temperatures using CFD software. The solver results are validated by comparing experimental results of Robinson and Hannemann’s standard force reference model HB-2. This research has successfully generated variation of skin temperature at $X$, $Y$, and $Z$ positions on external surface of cockpit at variable speed and angle of attack of aircraft along with governing equations of 2nd order polynomials. These governing equations will help the user to estimate skin temperature without rerunning simulation jobs in computational fluid dynamics design software. The results of a detailed analysis of the computational fluid dynamics design show that skin temperature under transient conditions at various angle of attack remains 0.9472 times the theoretical stagnation temperature with viscous heat dissipation.