Abstract: In order to obtain the analytical method to compute the circumferential strain on a soft pulse separation device (PSD), deformation processes of the middle section of the soft PSD, the medicine propellant grain and the case are simplified into a two-dimensional plane strain state. It is found that the main factors affecting the circumferential strain of the soft PSD are the circumferential strain of the inner surface of the propellant grain and the gap between the soft PSD and the propellant grain. In order to study the failure mechanism of the soft PSD in the double-pulse solid rocket motor (SRM), a two-dimensional axisymmetric finite element method (FEM) model of the stress process of the soft PSD is established. The variation of the strain of the soft PSD with the internal pressure load is obtained. It is found that the excessive circumferential strain is the main reason for the failure of the soft PSD. Comparing the analytical calculations with the FEM results, it can be found that the analytical method value is slightly higher than the FEM value, so the analytical method results can be used to initially estimate the circumferential strain of the soft PSD and then predict the rationality and feasibility of the design scheme. In order to further study the failure mechanism of the soft PSD, a micro-CT test of in situ stretching of the soft PSD material is carried out and the variation of porosity and elongation of the material is studied. The test results showed that when the material elongation is large, the microinterface debonding rapidly expands into a penetrating damage, and the PSD structure fails. The conclusions obtained in this paper can provide a useful reference for the design of double-pulse SRM. PubDate: Wed, 04 Dec 2019 13:05:03 +000
Abstract: In order to meet the needs of high-precision guidance for intercepting hypersonic targets, a novel head-pursuit guidance law considering the dynamic characteristics of a missile control system and the target mobility is presented via combining a fast power reaching law with backstepping sliding mode control in this paper. Initially, a three-dimensional head-pursuit system model of the missile and target is established. Subsequently, the system model is decomposed into a pitch plane system and lateral plane system, the control system dynamics are equivalent to second-order systems, and finite-time disturbance observers are introduced to estimate the target accelerations. On the basis of the previous work, the head-pursuit guidance laws of the vertical system and the lateral system which can stabilize the closed-loop system are designed separately and strict proofs of the methods are given. Finally, simulations are carried out to verify the effectiveness of this head-pursuit guidance law. PubDate: Tue, 03 Dec 2019 03:05:00 +000
Abstract: This paper studies the trajectory planning problem for multiple aircraft with logical constraints in disjunctive form which arise in modeling passage through waypoints, distance-based and time-based separation constraints, decision-making processes, conflict resolution policies, no-fly zones, or obstacle or storm avoidance. Enforcing separation between aircraft, passage through waypoints, and obstacle avoidance is especially demanding in terms of modeling efforts. Indeed, in general, separation constraints require the introduction of auxiliary integer variables in the model; for passage constraints, a multiphase optimal control approach is used, and for obstacle avoidance constraints, geometric approximations of the obstacles are introduced. Multiple phases increase model complexity, and the presence of integer variables in the model has the drawback of combinatorial complexity of the corresponding mixed-integer optimal control problem. In this paper, an embedding approach is employed to transform logical constraints in disjunctive form into inequality and equality constraints which involve only continuous auxiliary variables. In this way, the optimal control problem with logical constraints is converted into a smooth optimal control problem which is solved using traditional techniques, thereby reducing the computational complexity of finding the solution. The effectiveness of the approach is demonstrated through several numerical experiments by computing the optimal trajectories of multiple aircraft in converging and intersecting arrival routes with time-based separation constraints, distance-based separation constraints, and operational constraints. PubDate: Tue, 03 Dec 2019 02:05:01 +000
Abstract: The linear model of an aero engine is effective in a small range of the neighborhood of equilibrium points. According to this problem, the identification method for the parameter uncertain linear model of the aero engine was proposed. The identification problem is solved by calculating nonlinear programming. Considering the parameter uncertainty of the model is the critical point of this research during the optimization process. A parameter uncertain model of an aero engine can be obtained, which has large use range. This method is used for DGEN380 aero engine. The two parameters, VDD and VE, are defined for describing error range. Compared with experimental data, the uncertain model of DGEN 380 can simulate the real state of DGEN380 within 1% error range when . Compared with another conventional method of identification (recursive least squares), the parameter uncertain model, established by the method of this research, has a broad application area through parameter uncertainty of the model. PubDate: Mon, 02 Dec 2019 16:05:01 +000
Abstract: For the terminal guidance problem of a missile intercepting a maneuvering target, a profile-tracking-based adaptive guidance law is proposed with inherent continuity in this paper. To flexibly and quantitatively control the convergence rate of the line-of-sight rate, a standard tracking profile is designed where the convergence rate is analytically given. Then, a nonsingular fast terminal sliding-mode control approach is used to track the profile. By estimating the square of the upper bound of target maneuver, an adaptive term is constructed to compensate the maneuver. Therefore, no information of target acceleration is required in the derived law. Stability analysis shows that the tracking error can converge to a small neighborhood of zero in finite time. Furthermore, a guidance-command-conversion scheme is presented to convert the law into the one appropriate for endoatmospheric interceptions. Simulation results indicate that the proposed law is effective and outperforms existing guidance laws. PubDate: Sat, 30 Nov 2019 13:05:08 +000
Abstract: For the large-scale operations of unmanned aerial vehicle (UAV) swarm and the large number of UAVs, this paper proposes a two-layer task and resource assignment algorithm based on feature weight clustering. According to the numbers and types of task resources of each UAV and the distances between different UAVs, the UAV swarm is divided into multiple UAV clusters, and the large-scale allocation problem is transformed into several related small-scale problems. A two-layer task assignment algorithm based on the consensus-based bundle algorithm (CBBA) is proposed, and this algorithm uses different consensus rules between clusters and within clusters, which ensures that the UAV swarm gets a conflict-free task assignment solution in real time. The simulation results show that the algorithm can assign tasks effectively and efficiently when the number of UAVs and targets is large. PubDate: Tue, 26 Nov 2019 14:05:05 +000
Abstract: In order to simplify the manufacturing process or because of the limitation of the propulsion system, business jet, small civil airplane, and turboprop aircraft are always designed without leading-edge slats, which poses a great challenge to the flight safety during takeoff and landing. Focusing on the low-speed stall and poststall conditions, we investigated the aerodynamic characteristics and flow mechanism of high-lift configuration without slats using an improved delayed detached eddy simulation (IDDES) model which is validated by numerical simulations of the Common Research Model (CRM). Based on the analysis of the calculated results, conclusion can be made that the stall behavior of the configurations is directly related to the onset and evaluation of flow separation on the suction side. And through further research, an efficient evaluation method that is capable of qualitatively predicting the stall performance of two-element high-lift configuration by stall angle distribution of wing sections is proposed. By using the evaluation method, together with design rules summarized from the present study, high-lift configuration with mild-stall characteristic can be obtained in the preliminary stage of design. PubDate: Tue, 26 Nov 2019 11:05:08 +000
Abstract: In a bid to take advantage of natural characteristics of the proportional navigation guidance (PNG) in practical engineering, the PNG-based impact time control guidance (ITCG) continues to be a popular alternative for achieving the desired impact time of a missile. For most such ITCG, the performance is dependent on the accuracy of the time-to-go estimation. Along the lines of the development of PNG-based ITCG in earlier studies, a nonsingular ITCG is proposed on the basis of nonlinear formulations. It is demonstrated that, by theoretical analysis and numerical simulation, this proposed ITCG is shown to be advantageous in certain circumstances. By deriving a novel additional acceleration command, the proposed law is of lower dependence on time-to-go estimate and is capable of eliminating some singularities, leading to wider adjustable range of the desired impact time and better adaptability to more conditions. This research is expected to be supplementary to those presented in the current research literature. PubDate: Mon, 25 Nov 2019 03:05:02 +000
Abstract: Inflatable membrane reflectors are widely used in space and terrestrial deployable antennas. The mechanical properties of the inflated membrane reflector, however, are often a limiting factor in the level of the surface accuracy that can be achieved. In this study, membrane structural analysis is combined with sizing optimization to tailor the thickness of the membrane to improve surface accuracy of an axisymmetric parabolic inflated membrane reflector. Two reflector surface accuracy evaluation methods are employed in the optimizer and researched about their effect on the sensitivity analysis. Gradient-based optimization is employed together with a simplified sensitivity analysis, and the resulting algorithm is demonstrated on a membrane reflector case study. A linear thickness filter is used to avoid checkerboard patterns, and optimized solutions are computationally shown to achieve a feasible level of surface accuracy. PubDate: Sat, 23 Nov 2019 07:05:08 +000
Abstract: Aiming at the longitudinal motion model of the air-breathing hypersonic vehicles (AHVs) with parameter uncertainties, a new prescribed performance-based active disturbance rejection control (PP-ADRC) method was proposed. First, the AHV model was divided into a velocity subsystem and altitude system. To guarantee the reliability of the control law, the design process was based on the nonaffine form of the AHV model. Unlike the traditional prescribed performance control (PPC), which requires accurate initial tracking errors, by designing a new performance function that does not depend on the initial tracking error and can ensure the small overshoot convergence of the tracking error, the error convergence process can meet the desired dynamic and steady-state performance. Moreover, the designed controller combined with an active disturbance rejection control (ADRC) and extended state observer (ESO) further enhanced the disturbance rejection capability and robustness of the method. To avoid the differential expansion problem and effectively filter out the effects of input noise in the differential signals, a new tracking differentiator was proposed. Finally, the effectiveness of the proposed method was verified by comparative simulations. PubDate: Fri, 22 Nov 2019 08:05:16 +000
Abstract: In this paper, an observer-based adaptive backstepping attitude maneuver controller (briefly, OBABC) for flexible spacecraft is presented. First, an observer is constructed to estimate the flexible modal variables. Based on the proposed observer, a backstepping control law is presented for the case where the inertia matrix is known. Further, an adaptive law is developed to estimate the unknown parameters of the inertia matrix of the flexible spacecraft. By utilizing Lyapunov theory, the proposed OBABC law can guarantee the asymptotical convergence of the closed-loop system in the presence of the external disturbance, incorporating with the -gain performance criterion constraint. Simulation results show that the attitude maneuver can be achieved by the proposed observer-based adaptive backstepping attitude control law. PubDate: Thu, 21 Nov 2019 07:05:07 +000
Abstract: An adaptive saturated neural network (NN) controller is developed for 6 degree-of-freedom (6DOF) spacecraft tracking, and its hardware-in-the-loop experimental validation is tested on the ground-based test facility. To overcome the dynamics uncertainties and prevent the large control saturation caused by the large tracking error at the beginning operation, a saturated radial basis function neural network (RBFNN) is introduced in the controller design, where the approximate error is counteracted by an adaptive continuous robust term. In addition, an auxiliary dynamical system is employed to compensate for the control saturation. It is proved that the ultimate boundedness of the closed-loop system is achieved. Besides, the proposed controller is implemented into a testbed facility to show the final operational reliability via hardware-in-the-loop experiments, where the experimental scenario describes that the simulator is tracking a planar trajectory while synchronizing its attitude with the desired angle. Experimental results illustrate that the proposed controller ensures that the simulator can track a preassigned trajectory with robustness to unknown inertial parameters and disturbances. PubDate: Wed, 20 Nov 2019 10:05:09 +000
Abstract: The stability of a combustion chamber with a gas-centered swirl coaxial injector is investigated over a wide range of operating conditions in the aspect of thermoacoustic instabilities. First, flame shapes induced by the injector are analyzed for various recess lengths from experimental results. The spreading angle and flame size are reduced as the recess length increases. And, as a stability criterion, the damping factor is introduced. The operating conditions are divided into 7 sets with 9 tests, and fuel mass flow is fixed in each set. Stability boundaries are identified in terms of damping factor for the 63 tests by experiments without external perturbations. The stability map for the self-excited instability is obtained and it reveals two unstable regions. One is located at low dynamic head ratio and the other one is at high dynamic head ratio. A stable region is found at moderate dynamic head ratios. Relative stability is evaluated by various flow parameters over a wide range of operating conditions in the combustion chamber with a gas-centered swirl coaxial injector. PubDate: Tue, 19 Nov 2019 04:05:02 +000
Abstract: The second sharp-edged flight experiment is a faceted suborbital reentry body that enables low-cost in-flight reentry research. Its faceted thermal protection system consisting of only flat radiation-cooled thermal protection panels is cost-efficient since it saves dies, manpower, and storage. The ceramic sharp leading edge has a 1 mm nose radius in order to achieve good aerodynamic behaviour of the vehicle. The maximum temperature measured during flight was 867°C just before transmission ended and was predicted with an accuracy of the order of 10%. The acreage thermal protection system is set up by 3 mm fiber-reinforced ceramic panels isolated by a 27 mm alumina felt from the substructure. The panel gaps are sealed by a ceramic seal. Part of the thermal protection system is an additional transpiration-cooling experiment in which nitrogen is exhausted through a permeable ceramic matrix composite to form a coolant film on the panel. The efficiencies at the maximum heat flux are 58% on the porous sample and 42% and 30% downstream of the sample in the wake. The transient load at each panel location is derived from the trajectory by oblique shock equations and subsequent use of a heat balance for both cooled and uncooled structures. The comparison to the heat balance HEATS reveals heat sinks in the attachment system while the concurrence with the measurement is good with only 8% deviation for the acreage thermal protection system. Aerodynamic control surfaces, i.e., canards, have been designed and made from a hybrid titanium and ceramic matrix composite structure. PubDate: Mon, 18 Nov 2019 09:05:07 +000
Abstract: The load characteristics for the flow regulator of a solid ducted rocket are discussed in this paper. The mechanism and the influence factor of the load in the flow regulator were studied both theoretically and experimentally, and the system load and the working load were divided according to the mechanism. Additionally, the load influence on the working quality of the flow regulator and the ducted rocket were analyzed. System modeling of the flow regulator was carried out based on the working mechanism of the gas generator and the Stribeck friction model, and models of the actuator, gas generator, and load were built. Furthermore, considering the identification of the model parameters and the inapplicability of the conventional Stribeck identification method to the flow regulator, a new Stribeck model parameter identification method was proposed. A sine wave was used as the signal source, and specially designed devices were used to accomplish the switch of the working condition. Through the use of the genetic algorithm, the rotational inertia and parameters of the system load and working load were identified by stages. Finally, the validation of the identification method was carried out with a simulation. PubDate: Mon, 18 Nov 2019 08:05:07 +000
Abstract: A morphing trailing-edge (TE) wing is an important morphing mode in aircraft design. In order to explore the static aeroelastic characteristics of a morphing TE wing, an efficient and feasible method for static aeroelastic analysis has been developed in this paper. A geometrically exact vortex lattice method (VLM) is applied to calculate the aerodynamic forces. Firstly, a typical model of a morphing TE wing is chosen and built which has an active morphing trailing edge driven by a piezoelectric patch. Then, the paper carries out the static aeroelastic analysis of the morphing TE wing and corresponding simulations were carried out. Finally, the analysis results are compared with those of a traditional wing with a rigid trailing edge using the traditional linearized VLM. The results indicate that the geometrically exact VLM can better describe the aerodynamic nonlinearity of a morphing TE wing in consideration of geometrical deformation in aeroelastic analysis. Moreover, out of consideration of the angle of attack, the deflection angle of the trailing edge, among others, the wing system does not show divergence but bifurcation. Consequently, the aeroelastic analysis method proposed in this paper is more applicable to the analysis and design of a morphing TE wing. PubDate: Sat, 16 Nov 2019 12:05:08 +000
Abstract: Folding fins are widely adopted in missiles for the efficient use of space during storage and transportation, while nonlinear behavior of freeplay is inevitable due to the factors such as mismachining tolerance, assembly error, and abrasion. The problem of nonlinear system identification of folding fins with freeplay is considered in this paper. A direct parameter estimation method which can identify the nonlinear system with freeplay under base excitation is proposed and subsequently applied to establish the nonlinear dynamic model of a folding fin. The best set of coefficients is selected by using the significance test, allowing the proposed method to detect and locate the most relevant nonlinearities of the practical structure. Experimental results demonstrate that the proposed method is able to decouple the linear and nonlinear dynamics of a nonlinear structure and estimate natural frequencies of the derived linear system along with nonlinear internal forces in one computational step, even if no a priori knowledge of the type of nonlinearities is given. PubDate: Fri, 15 Nov 2019 10:05:09 +000
Abstract: This paper focuses on the aerodynamics and design of an unmanned aerial vehicle (UAV) based on solar cells as a main power source. The procedure includes three phases: the conceptual design, preliminary design, and a computational fluid dynamics analysis of the vehicle. One of the main disadvantages of an electric UAV is the flight time; in this sense, the challenge is to create an aerodynamic design that can increase the endurance of the UAV. In this research, the flight mission starts with the attempt of the vehicle design to get at the maximum altitude; then, the UAV starts to glide and battery charge recovery is achieved due to the solar cells. A conceptual design is used, and the aerodynamic analysis is focused on a UAV as a gliding vehicle, with the calculations starting with the estimation of weight and aerodynamics and finishing this stage with the best glide angle. In fact, the aerodynamic analysis is obtained for a preliminary design; this step involves the wing, fuselage, and empennage of the UAV. In order to achieve the preliminary design, an estimation of aerodynamic coefficients, along with computational fluid dynamics analysis, is performed. PubDate: Thu, 14 Nov 2019 14:05:08 +000
Abstract: This paper studies the effect of continuous gusts on the initial disturbance of the airborne missile’s horizontal backward derailment on large transport aircraft. The longitudinal fluctuation of the airborne launching platform under continuous gust was obtained with different calculation methods, and the finite element model of the derailing process was established then verified by experiments. Finally, combined with the longitudinal fluctuation of the launching platform and the finite element model of the missile and launching platform, the influence of platform fluctuation on separation time, missile speed, pitch angle, and descending distance while derailing was studied and analyzed. It is found that the longitudinal fluctuation of the launching platform is similar to that of the gust but lags behind, and the missiles under the fluctuating platform have longer separation time, lower derailing speed, and greater derailing pitch angle and descending distance. PubDate: Tue, 12 Nov 2019 16:05:04 +000
Abstract: The fly-by-wire system plays an important role in modern civil aircraft. As a typical safety-critical system, its reliability will affect the safety of aircraft significantly. In the paper, stochastic Petri nets are applied in the reliability modeling and analysis for the fly-by-wire system to represent its dynamic (time-dependent) failure behaviors. Stochastic Petri net-based reliability models are established for all kinds of architectures including series, parallel, -out-of-, warm standby, cold standby, and load-sharing architectures, which are commonly used in the fly-by-wire system. A Monte Carlo simulation method is proposed for the stochastic Petri net-based reliability models to generate system lifetime samples, and the system reliability parameters can be calculated in terms of the lifetime samples. Finally, a fly-by-wire system is used as a case study to illustrate the application and effectiveness of our proposed approaches. The results show that the error of the reliability value in a flight duration obtained by our Monte Carlo simulation method is less than compared with the analytical equation. PubDate: Tue, 12 Nov 2019 16:05:02 +000
Abstract: Eu:CROPIS is DLR’s first mission of the Compact Satellite Program. Its primary payload focuses on the research of closed-loop biological, regenerative life support systems, in a simulated gravitational environment of the Moon and Mars over months at a time. This is achieved by rotation of the satellite around its central body axis, using only magnetic torquers as actuators. A secondary payload (“PowerCells”) by the NASA Ames Research Center also utilizes the artificial gravity to conduct growth experiments on genetically modified organisms (GMOs). These payloads and the system design imposed constraints which affected the Assembly Integration and Verification (AIV) program in various ways and created challenges for the relatively small team to find solutions for. The paper to be presented will address the different aspects of the AIV program. This includes the verification of different critical components like the newly developed CFRP pressure vessel containing the primary payload and the Micrometeoroid and Debris Protection Shield, which protects it. Both items went through rigorous testing, including high-velocity impact tests, to ensure their reliability in orbit. Various other aspects concerning the biology had to be taken into account during AIV campaigns: due to the presence of degradable components within the primary payload, a late access capability had to be implemented in order to exchange biology as well as chemistry in cases of launch delays. To allow these operations as close as six months prior to launch, a highly flexible and streamlined acceptance test campaign was developed. A major impact on test planning and logistics was the fact that the secondary payload “PowerCells” contains GMOs, which European and German regulations restrict to be handled exclusively in especially certified laboratories (biosafety level 1 (BSL-1)). Thus, the use of external test facilities for the flight model campaign was not feasible as no European test center is certified to BSL-1. In consequence, the clean room facilities of the DLR Institute of Space Systems had to be certified to BSL-1 and new test infrastructure had to be procured in a short time frame to cover for acceptance testing. The design of the satellite and nature of the attitude control subsystem required limits on the magnetic momentum of the system and every unit it contains. A test flow incorporating the magnetic property measurement of each unit and a final system-level test in an external facility had to be devised, which enabled budgeting and projection of expected measurement results on the system level. Furthermore, the moments of inertia had to be measured precisely in order to have a stable spinning axis enabling a stable gravity simulation. Finally, the functionality had to be verified for each unit and for the system which required that several small test campaigns had to be conducted, like a solar panel deployment test and extensive software testing. A tight link to the operations teams of the German Space Operations and Control Center during such tests and beyond finally ensures the operability of the overall system in the operational phase. PubDate: Mon, 11 Nov 2019 10:05:18 +000
Abstract: A search and rescue mission is a typical UAV network application scenario. In this case, it is necessary to deliver messages quickly and efficiently to the ground station through mutual cooperation between UAVs. Many methods used in this case have problems such as unbalanced popularity (the ratio of the relayed message number to the total message number) of nodes, large proportion of ping-pong effect, and long delay. In view of the above problems, this paper proposes a method named PB (Popularity Balance Method for UAVs in the same area layer) based on division of the whole search area. The method divides the search area into multiple area layers. Message transmission between area layers adopts a geographical routing manner, that is, messages are transmitted to the area layer closer to the ground station. The division of the search area changes the pattern of message transmission. Messages are delivered to the area layer closer to the destination node rather than the node closer to the destination. The pattern causes messages’ passing direction to be replaced by “point-to-point closer” to “point-to-face closer.” On the basis of message transmission at area layers, reasonable planning of UAVs’ distribution can effectively improve the network performance deterioration caused by a “hot spot.” Both analysis and experiments show that PB is superior to some existing methods in popularity balance of nodes and ping-pong effect. In addition, experiments also show that it gets better results in targets of delay, delivery rate, and hop count. PubDate: Mon, 11 Nov 2019 00:07:37 +000
Abstract: The inflatable buffer landing system, such as the airbag, has been widely used for deep space exploration as a low-cost entry. The impact performance of an airbag landing attenuation system on Mars had to be proven through cushion testing on Earth. In this paper, a design method for the airbag landing attenuation system on Earth is proposed based on similarity relation. With this method, the impact response of the spherical airbag landing attenuation system is derived based on assumptions, and the principal factors have great influence on the impact response of the landing system have been proved. Then, the similarity relation between the full-scale airbag landing attenuation system in the Martian atmosphere and its Earth prototype is obtained through theoretical derivation. Finally, the proposed similarity relationship has been tested on several ground prototypes by FEA. The results of the study show that the Earth prototype by the developed design method is capable of predicting the impact response with good accuracy. PubDate: Mon, 11 Nov 2019 00:07:34 +000
Abstract: This paper presents a study of the most suitable ground test equipment for the Airbus A400M aircraft aimed at optimizing a design that combines portability, flexibility, and configurability. The study is based on current trends in the configuration of the different parts that are involved in the process of designing Aircraft Interface Modules (AIMs). Researches have been conducted facing the real problems of data acquisition and control in the assembly, maintenance, and repair processes of aircraft. In this sense, the number of signals used simultaneously in historical tests has been determined to be a parameter that plays an important role in the analysis. This work is within the framework of a project that our working group of the Department of Electronic Technology has carried out together with the company Airbus Defense and Space (Airbus DS). Thus, all the analyses and researches for the different configurations of the test equipment have been validated for the company. Some state-of-the-art systems have been considered in the process of designing the structure of the test system while analyzing the main advantages and drawbacks. Every hardware or software component was identified and justified according to its importance with respect to the use case. A comparison of how every approach meets the requirements was taken into account to define a methodology for designing a portable and configurable ground testing system. Thus, the conclusions of our study are being used by Airbus DS for the design of a prototype for the next generation of AIMs in A400M. PubDate: Mon, 11 Nov 2019 00:07:31 +000
Abstract: The loss of fan blades in an aeroengine, or fan-blade out (FBO), is a type of accident that causes a sudden imbalance and large impact load, which leads to complex vibration of a system. To conduct a dynamic analysis of an aeroengine rotor system is an important requirement for relevant departments. The purpose of this paper is to study the dynamic response of a complex dual-rotor system suffering FBO events and the protective effect of the fusing structure on the system. The dynamic model of an aeroengine dual-rotor system is established, and the response of the rotor system is obtained by calculation and analysis. The rear support bearing of the fan has a high reaction force, which may lead to bearing failure. The mechanism of a fusing structure is analyzed, and the results show that the sudden imbalance will produce impact loads on the rotor, resulting in a sharp increase in the vibration amplitude and reaction force, and then, attenuation to steady state. The fusing structure can reduce the amplitude of steady-state rotor vibration and reaction force on the support bearings. However, the transient response of the rotor will increase because of the sudden change in support stiffness. PubDate: Mon, 04 Nov 2019 03:05:03 +000
Abstract: A new low-thrust trajectory design method is proposed that is based on the finite Fourier series method with pseudoequinoctial elements rather than the more common cylindrical coordinate components. The bijection relation between the elements and control variables is ensured by introducing an additional equality constraint derived from the angular momentum conservation. The guidance law and on-line control variables are obtained by applying inverse dynamics and the framework of inverse simulation technology, respectively. The pseudoequinoctial finite Fourier series method has the advantages of both the Fourier series and the perturbation analysis methods. For two-body problems, three cases were studied: the Earth to Mars, 1989ML, and Tempel-1 missions. Regarding the design of a rendezvous trajectory with a large inclination angle and a high eccentricity rate, this method yields a broader range of feasible results than the traditional Fourier series method. The circular restricted three-body problem was solved for the first time using the pseudoequinoctial finite Fourier series method combined with the patched conics method. The low-thrust Earth-Moon transfer was analyzed, and the results show that this method improves window analysis efficiency and guarantees precision of the initial geocentric trajectory for the low-thrust transfer. PubDate: Mon, 04 Nov 2019 03:05:02 +000
Abstract: For small unguided space interceptors, the interception probability is an important index to evaluate their strike capability. However, the Monte Carlo (MC) based simulation method not only consumes a lot of computation time but also theoretically lacks interpretability. In this regard, this paper proposes an analytical method that can quickly and accurately calculate the short-term space interception probability. Firstly, by considering the effect of perturbation force on spacecraft as the effect of external force acceleration, the analytical calculation formula of the short-term state error covariance propagation of space target and interceptor is deduced. Next, by projecting the state error and rotating the coordinate system, the joint error distribution of the target and the interceptor in the calculation coordinate system at the time of closest approach (TCA) is obtained. Thereby convert the calculation of the space interception probability into the integral of the 2-dimensional probability density function in the circular domain. Then, the Laplace transform and Taylor expansion are used to obtain the exact power series expression and the maximum truncation error of the integral calculation, and the analytical calculation of the short-term space interception probability is realized. Finally, the effectiveness of the proposed method is verified by a simulation example. The proposed method can directly calculate the space interception probability according to the initial state error distribution of the target and interceptor, and the whole calculation process does not contain double integral operation. The proposed method has high computational efficiency, is suitable for on-orbit calculation, and provides effective support for the rapid evaluation of the strike capability of the space interceptor. PubDate: Sun, 03 Nov 2019 00:07:08 +000
Abstract: In aerospace engineering, structural modifications play an essential role in design of structures. In some cases, it is necessary to guarantee that a specific natural frequency of the structure remains unchanged when additional masses are attached. The methods based on the Sherman-Morrison formula are proposed in this paper, called the optimal selection method and the absolute value method, to maintain the specific natural frequency. The methods are both implemented by installing a spring on the system and can eliminate the effect of the additional mass on the specific frequency. The proposed methods were verified to be effective and accurate through numerical simulations. Results show that the optimal selection method has similar applicability as the existing real value method, and both methods are applicable only in cases of small damping. In addition, the absolute value method has extensive applicability in systems with either small or large damping. PubDate: Sun, 03 Nov 2019 00:07:06 +000
Abstract: A self-designed oxygen-kerosene ablation system was employed to research the ablation properties of silicone rubber-based insulation material under different oxygen-rich conditions, that is, 0%, 5.00%, 7.71%, 17.01%, and 18.50%. The morphology of posttest specimens was analyzed via a scanning electron microscope (SEM), and the ablation rates were calculated. Experimental results showed that when the particle concentration was a fixed value, the mass and the linear ablation rates increased first and then decreased with the rise of oxygen content; the maximum values were 0.572 g/s and 0.933 mm/s, respectively. Under high oxygen-rich conditions, the formation of silicone rubber restricted the further increase of the ablation rates, filling more gaps of the reaction layer with liquid silicone rubber. Meanwhile, the thickness of the liquid glass layer attached to the surface was increased, which acted as a buffer against the impacts and erosion of particles and retarded the consumption and recession of materials. Excess oxygen spread to the edge of the ablation pit and reacted. Finally, these led to the increase in diameter of the ablation pit and the decrease of two types of ablation rates. PubDate: Wed, 30 Oct 2019 10:05:14 +000
Open Access journal
[338 journals]
