Journal of Aerospace Engineering

As one of the most promising propulsive systems for future hypersonic vehicles, the hydrogen-fueled dual-mode scramjet combustor has drawn the attention of an ever increasing number of researchers. The two-dimensional coupled implicit NS equations, the standard $k-\epsilon$ turbulence model, and the finite-rate/eddy-dissipation reaction model have been applied to numerically simulate the flow field of the dual-mode scramjet combustor, including the conditions of engine ignition and cold flow. The effect of the injection pressure and temperature on mode transition, and the movement of the shock-wave train, has been discussed. At the same time, the influence of the injection modes, namely, the transverse and horizontal, on the flow field of the combustor has been investigated. The obtained results show that the mode transition can be carried out by increasing the injection pressure and decreasing the injection temperature simultaneously. However, when the injection pressure is too high, the shock-wave train is pushed out of the isolator toward the entrance, and this causes inlet unstart. It has been found that the effect of the injection temperature on mode transition is smaller than that of the injection pressure. It is not easy to achieve mode transition when the fuel is injected horizontally into the airflow, and the combustion efficiency is lower than when the fuel is injected transversely.

An expander cycle is one potential fuel feeding cycle for a scramjet, which has been successfully applied in the liquid rocket. However, very little literature exists about the fuel feeding cycle for a scramjet. A recooling cycle has been newly proposed for a scramjet to reduce the fuel flow for cooling, the secondary function of which acts as the fuel feeding cycle. To illustrate the difference between the expander cycle and the recooling cycle, the model of the components, the pump, the turbine, and the cooling-channel are proposed in terms of hydrodynamic, thermal, power balance, and Machnumber constraints. The applicability of expander cycle as the fuel feeding cycle for a scramjet is investigated, and then the differences in working principle, functions, components, performance, and so on between the expander cycle and the recooling cycle are discussed. Furthermore, the new limitations introduced by the recooling cycle are also analyzed.

The application of fuel cell (FC) technology to aircraft propulsion and/or energy supply is becoming of great interest for undoubted advantages in terms of pollution emissions and noise reduction. A better understanding of problems related to fuel cells applied to aeronautics is sought by the European Commission (EC) funded project Environmentally Friendly Inter-City Aircraft Powered by Fuel Cells (ENFICA-FC). The main objective of the ENFICA-FC project was to develop and validate the use of a fuel cell–based power system for the propulsion of more-electric/all-electric aircraft. The fuel cell system was installed in the light sport aircraft Rapid 200, which was flight and performance tested. One of the key items under investigation is the simulation of the cooling system and the evaluation of fuel cell temperature. The polymer electrolyte membrane fuel cell (PEMFC) is considered to be the best candidate for the fuel cell vehicle because it has high power density, solid membrane electrolyte, and as it operates at low temperatures, it has a fast start-up. However, to generate a reliable and efficient power response and to prevent membrane degradation or damage with hydrogen and oxygen depletion, a sophisticated control technique becomes vitally important. In particular, as the ionic conduction of the polymeric membrane is a function of its degree of humidification, the stack temperature has to be carefully controlled to avoid phenomena of water evaporation, causing an increase of ohmic drop and a decrease of stack performances. The output voltage and hence the power of the fuel cell system is affected considerably by the change of the stack temperature. A simplified fluid-dynamic model has been developed and validated by computational fluid dynamics (CFD) analysis and it is used to compute the air flow to the fuel cell heat-exchanger inlet. Propeller effects are included referring to an optimal propeller specifically designed for the ENFICA-FC project. A mathematical model of the fuel cell system dynamics coupled with the fluid-dynamic model was studied in detail and experimentally validated during two flight tests of the Rapid 200-FC.

This paper reviews recent advances and explores new trends in designing configuration and materials for application in very light jet (VLJ) aircrafts. Current technical information of the light jet aircraft in the market have been studied for competitor analysis purpose; the review of market design considers maximum take-off weight (MTOW), empty weight (WE), maximum cruise speed ($V_{\mathrm{cr}·\max }$), direct operating cost (DOC), and aircraft price. Recent trends in aircraft configurations have been reviewed with particular reference to both conventional and unconventional designs. The existing technology of lightweight aerospace materials has been reviewed and the recommendations have been made for applications of composite and appropriate advanced materials in new VLJ aircrafts.

High–surface area micro-and nanoscale electrospun fibers were produced for noise-reduction purposes using three different polymers: polyvinylpyrrolidone (PVP), polystyrene (PS), and PVC. These polymers were initially dissolved in appropriate solvents and electrospun at various electrospinning conditions. The two-microphone transfer-function method of the Brüel and Kjær (B&K) impedance tube was used to determine the acoustical properties of the electrospun fibers at different frequencies. Test results showed that the absorption coefficients of the fibers (200 nm to 7 µm) were significantly enhanced at nanoscale. The reason behind this phenomenon may be attributed to the higher surface area of fibers and their interactions with more sound waves/air molecules. Overall, the findings of this research will open up new possibilities for the noise-reduction problem of aircraft, other transportation vehicles, and infrastructures.

A macrolevel finite element–based model has been developed to simulate the mechanical and impact response of triaxially braided polymer matrix composites. In the analytical model, the triaxial-braid architecture is simulated by using four parallel shell elements, each of which is modeled as a laminated composite. For the current analytical approach, each shell element is considered to be a smeared homogeneous material. The commercial transient dynamic finite-element code LS-DYNA is used to conduct the simulations, and a continuum damage mechanics model internal to LS-DYNA is used as the material constitutive model. The constitutive model requires stiffness and strength properties of an equivalent unidirectional composite. Simplified micromechanics methods are used to determine the equivalent stiffness properties, and results from coupon-level tests on the braided composite are utilized to back out the required strength properties. Simulations of quasi-static coupon tests of several representative braided composites are conducted to demonstrate the correlation of the model. Impact simulations of a represented braided composite are conducted to demonstrate the capability of the model to predict the penetration velocity and damage patterns obtained experimentally.

A piecewise linear parametric uniaxial stress-strain approach has been used to obtain the closed form nonlinear moment curvature response on the basis of strain compatibility in bending for epoxy resin materials. The stress-strain curves, consisting of a bilinear ascending curve followed by strain softening and constant plastic flow in tension and compression, are described by two main parameters, with an additional five nondimensional tensile and seven nondimensional compressive parameters. The main parameters are the modulus of elasticity and strain at the proportional elastic limit point in tension. Parametric studies show that ultimate tensile stress and compressive yield stresses and tension and compression flow stresses have the highest effects on flexural load carrying capacity. Moment curvature equations, in conjunction with softening localization and static equilibrium conditions, were used to simulate the flexural load-deflection response of a beam under three-point bending (3PB) conditions. The simulations and experiments reveal that the direct use of uniaxial tensile and compressive stress-strain curves underestimates the flexural response caused by the stress gradient and the differences in the effective volume of the body subjected to critical stresses.

Impact response and damage induced by a drop-weight instrument on glass reinforced (GLARE) 5 fiber-metal laminates of different thicknesses were studied. The effect of the impactor mass was also considered. The damage characteristics were evaluated using both the nondestructive ultrasonic and mechanical-sectioning techniques. Only the contour of the entire damage area can be obtained using ultrasonic C-scan; whereas, more details of the damage are provided through the mechanical cross-sectioning technique. Thicker GLARE 5 offers higher impact resistance. It was found that the failure mode changed with varying thickness and impactor mass. Panels that were tested with the lighter impactor mass would reach the perforation limit faster than the panels tested with the heavier impactor mass for the same impact energy. It was observed that the critical behaviors of the fiber and aluminum depended on the thickness of the panel. Threshold cracking energy varied parabolically with respect to the specimen thickness. Thus, impact energy alone cannot be used as a scale parameter of damage size.

The rotational motion of a heavy solid about a fixed point in the presence of a gyrostatic moment vector is investigated in this paper. It is supposed that the body is rapidly spinning about the major or the minor principle axis of the ellipsoid of inertia. The Krylov-Bogoliubov-Mitropolski technique is modified to obtain the periodic solutions of the equations of motion of the body with nonzero basic amplitude. These solutions are performed by computer codes to get their graphical representations. The result of this study was compared with similar previous works.

It is very difficult to avoid the orthoerror and gain error in the three-axis measuring device by means of improving the manufacturing process, and the measure precision is tremendously limited by these errors. A model for this study was built by the mapping relationship between the measure system and the virtual orthosystem. A genetic algorithm was proposed for calibrating the orthoerror and gain error among three measuring axes and was compared with other existing approaches presented in the literature. The results of simulation demonstrate that the genetic algorithm shows a better performance than other algorithms, and a triaxial measure system can be calibrated more efficiently by genetic algorithm.

An integrated approach combining a state-dependent Ricatti equation (SDRE) and an extended Kalman filter (EKF) is applied to spacecraft formation flying for robust orbital and attitude maneuvers. The formation flying considered in this study is a two-spacecraft formation with a bounded out-of-orbit plane relative motion. The chaser is required to simultaneously perform large position and angle maneuvers with sufficient accuracy. The chaser is then required to maintain a relative orbit expressed in the local-vertical-local-horizontal frame with respect to the target and align its attitude with the target attitude for more than one orbital period. The controls of the chaser are formulated as a nonlinear optimal regulator problem using their highly nonlinear dynamics. To test the robustness of this integrated approach, highly nonlinear dynamics, including external disturbances, are employed as the truth plant that is then used to generate the measurements for an EKF. The tracking error is then computed between the true plant and the desired state. A six degrees-of-freedom simulation of a two-spacecraft flying formation is used to demonstrate the robustness of this integrated approach to external disturbances and plant uncertainties. The integrated approach also has an effect well under the large uncertainty of the chaser moment of inertia.

In this paper, linear-quadratic regulator (LQR) method is proposed for the design of nonlinear spacecraft control systems. The proposed design is based on the linearized spacecraft model that involves only three components of the quaternion. By using a simple and special structure of the linearized reduced quaternion model, an analytical solution for the controller. It is shown that the analytical solution of the state feedback matrix is an explicit function of the cost matrices $Q$ and $R$. The analytic solution makes it convenient in design practice to tune the feedback matrix and the cost matrices $Q$ and $R$ to balance the requirements between response performance and fuel consumption. It is shown that the designed controller globally stabilizes the nonlinear spacecraft system, whereas it locally optimizes the spacecraft performance. A design example is provided to show the effectiveness of the design method.

Multifunctional power structures (MFPS) are fully integrated subassemblies that perform both structural and power functions for spacecraft. By combining functions across subsystems into single units, mass and volume savings can be achieved. Focusing on battery-based MFPS in Earth-orbiting spacecraft, the embedded lithium ion batteries that are used have strict temperature limits, outside of which efficiency and safety is compromised. Considering the limits of the model’s prediction accuracy, numerical simulation has shown that a range of Earth orbits exist where an MFPS mounted in a deployed solar array would not require the addition of further thermal control with respect to the nonmultifunctional case. The numerical simulation consisted of a lumped parameter reduction of the model to a discrete set of layers. Thermal control is required to prevent overcooling of the battery in eclipse and to extend the range of orbits where MFPS can be used. An assessment of current thermal control was performed to establish the viability of each technology, with viability defined by feasibility and the mass of the system. The use of coatings, insulation, heaters, and phase-change materials were considered. It was found that the range of viable orbits is dependent on the quantity of MFPS savings that can be sacrificed.

Tongji-1 lunar soil simulant (TJ-1 for short) was recently invented by Tongji University, China. It could be used to construct testing ground in the preparatory impact tests of lunar landers under the extremely adverse condition in the Earth environment, where the interaction pressure between lunar lander and lunar soil has the important effect on the safety of lunar lander. The TJ-1 is originated from red volcanic ash deposits collected from northern China. The raw materials are firstly dried, crushed, and sieved to obtain materials of seven particle sizes. Gradation is then used as the controlling parameter to generate TJ-1 with desired mechanical properties. The TJ-1 demonstrates certain similarity to lunar regolith in physical and mechanical properties, and it is an ideal substitute of lunar regolith located from 0 to 30 cm below lunar surface in the mare region on the Moon. Compared with other simulants introduced in the literature, TJ-1 has an internal frictional angle as high as 45° at $e=1.0$ and hence is advantageous over other simulants when high shear strength is desired. The TJ-1 will be a good choice as a ground material in the preparatory impact tests of lunar landers in the Chinese Lunar Exploration Program.

The removal of lunar dust adhered to an astronaut’s spacesuit is critical for long-term lunar exploration. The author is developing a cleaning device that utilizes electrostatic force; this system employs a combination of electrostatic capture and electrostatic transport of dust. High voltage is applied between a screen electrode and an aluminum film deposited on a Mylar sheet positioned under the surface fabric of the spacesuit. By means of electrostatic force, dust adhered to the fabric is captured at the screen electrode of the device. The captured dust is then transported by an electrostatic traveling wave and transferred to a collection bag. The capture rate increased with the amount of dust adhered to the fabric and reached a value of 82%. Although small particles trapped between fabric fibers were difficult to remove, this system should be used in preliminary cleaning of spacesuits, saving valuable time for astronauts on the Moon.