Search

This paper addresses the scheduling problem arising from refueling multiple geosynchronous (GEO) satellites with multiple servicing vehicles [many-to-many (M2M) refueling]. The problem is defined by a set of potential fuel tanker locations, a homogeneous fleet of servicing vehicles with limited capacities, and a set of fuel-deficient GEO satellites with known fuel demands. The objective is to open a subset of fuel tankers, to assign GEO satellites to these tankers and to design vehicle servicing sequences, in order to minimize the total mission costs. To achieve such an economical refueling strategy, the fuel tanker location and routing decision are required to be determined simultaneously. It is shown that this problem can be formulated as a location-routing problem (LRP) which is obviously NP-hard. In order to solve it, an ant colony optimization (ACO) metaheuristic that features a special solution representation scheme is proposed. The proposed algorithm is then tested on two types of instances depending on whether the cost for opening a fuel tanker is considered or not. Numerical experiments show that both the consideration of the fuel tanker opening costs and their locations do affect the refueling schedule. Furthermore, the proposed algorithm outperforms the existing approaches in the M2M refueling mission design, and it is able to detect the optimal refueling strategy on all instances containing 85 GEO targets with 15 potential fuel tankers while previous works stopped at 15 satellites and 2 fuel tankers. • Scheduling of the multiple satellites refueling mission is presented. • Servicing vehicle routing and the fuel tanker location are considered simultaneously. • An ant colony optimization based solving methodology is proposed. • A special solution presentation scheme is applied to improve the computational burden. [ABSTRACT FROM AUTHOR]

Abstract The ninth edition of the China Trajectory Optimization Competition (CTOC9) considered the problem of navigation augmentation over a set of Chinese cities, wherein the competitors were expected to design a suitable LEO constellation to obtain the required navigation performance with the aim of minimizing overall system cost. In addition, an economical deployment strategy was also considered. The synthesis methods used by the Xi'an Satellite Control Center (XSCC) are described, along with the promising solution found by the team. Firstly, two techniques, feasible region based reduction and coverage region based reduction, are applied in sequence to narrow the search space. Then, the optimal constellation configuration is obtained through an Differential Evolution Algorithm (DE). Finally, a series of maneuver strategies are searched to inject each navigation satellite from a given initial carrying orbit to its target location. The solution obtained during the competition ranked third, and the improved one, found after the deadline, got the outstanding prize. [ABSTRACT FROM AUTHOR]

Parametric modeling of aircrafts plays a crucial role in the aerodynamic design process. Effective parametric approaches have large design space with a few variables. Parametric methods that commonly used nowadays are summarized in this paper, and their principles have been introduced briefly. Two-dimensional parametric methods include B-Spline method, Class/Shape function transformation method, Parametric Section method, Hicks-Henne method and Singular Value Decomposition method, and all of them have wide application in the design of the airfoil. This survey made a comparison among them to find out their abilities in the design of the airfoil, and the results show that the Singular Value Decomposition method has the best parametric accuracy. The development of three-dimensional parametric methods is limited, and the most popular one is the Free-form deformation method. Those methods extended from two-dimensional parametric methods have promising prospect in aircraft modeling. Since different parametric methods differ in their characteristics, real design process needs flexible choice among them to adapt to subsequent optimization procedure. [ABSTRACT FROM AUTHOR]

Hypersonic vehicles have been paid more and more attention by various countries, and many developed countries have set off a competition of hypersonic technology. Under this background, this paper creatively applies the reconstruction method based on three vanishing points to the reconstruction of complex aircraft for the first time. The shape of aircraft is restored by photos from different angles, which provides a good help for preliminarily understanding and mastering the information of international mainstream advanced aircraft. In this paper, the feasibility of the whole method is verified by the classical X-37B model, and its three-dimensional structure model can be obtained relatively accurately. In terms of aerodynamic performance evaluation, the panel method is adopted to evaluate the aerodynamic performance on the basis of verifying the reliability of the engineering method. And the results show that the lift-to-drag ratio reaches the maximum of 1.22 at the angle of attack of 25°. This is quite similar to the CFD calculation of class X-37B model in other literatures. The accuracy of the model makes up for the error of the engineering estimation method to a certain extent. In view of the inaccuracy of the wing reconstruction, the genetic algorithm is used to optimize its aerodynamic performances. The lift-to-drag ratio of the wing is increased by 26.3% at 25°, which leads to an increase of 4.35% in the optimal angle of attack lift-to-drag ratio of the whole model. At last, the modularization programming approach is adopted in the optimization process, ensuring the generality of this method, and the whole method will have a broad application prospect. • A reusable vehicle was reconstructed by the developed reverse engineering approach. • The aerodynamic performance of the reusable vehicle was optimized. • When the angle of attack is 25°, the lift-to-drag ratio reaches the maximum value of 1.22. • The overall lift-to-drag ratio of the model is with an increase of 4.35%. [ABSTRACT FROM AUTHOR]

The mixing and combustion process has an important impact on the engineering realization of the scramjet engine. The nonreacting and reacting flow fields in a transverse injection channel have been investigated numerically, and the predicted results have been compared with the available experimental data in the open literature, the wall pressure distributions, the separation length, as well as the penetration height. Further, the influences of the molecular weight of the fuel and the jet-to-crossflow pressure ratio on the wall pressure distribution have been studied. The obtained results show that the predicted results show reasonable agreement with the experimental data, and the variable trends of the penetration height and the separation distance are almost the same as those obtained in the experiment. The vapor pressure model is suitable to fit the relationship between the penetration height, the separation distance and the jet-to-crossflow pressure ratio. The combustion process mainly occurs upstream of the injection port, and it makes a great difference to the wall pressure distribution upstream of the injection port, especially when the jet-to-crossflow pressure ratio is large enough, namely 17.72 and 25.15 in the range considered in the current study. For hydrogen, the combustion downstream of the injection port occurs more intensively, and this may be induced by its smaller molecular weight. [ABSTRACT FROM AUTHOR]

The wide-speed range characteristics of aerodynamic configuration for aircraft are critical to the design of the hypersonic vehicle, and many novel strategies have been proposed. As one of the valid schemes of aircraft with a wide-speed range, the vortex lift waverider with a wide-speed range has attracted an increasing attention. In the current study, the low speed aerodynamic performance of vortex lift waveriders with a wide-speed range has been investigated numerically by means of the three-dimensional implicit Reynolds-averaged Navier-Stokes (RANS) equations coupled with the SST k-ω turbulence model. At the same time, the numerical approaches employed have been validated against the available experimental data in the open literature. The obtained results show that the low speed lift-to-drag ratio of the cuspidal waverider is higher than that of the general osculating cone waverider, while the low speed lift-to-drag ratio of the delta-winged waverider is lower than that of the general osculating cone waverider except for the small and negative angles of attack. At the same time, the low speed vortex lift characteristics of the three types of osculating cone waveriders have been compares and analyzes, and some valuable conclusions have been obtained. • Low speed aerodynamic performances of vortex lift waveriders were analyzed and compared. • Leading edge vortex lift is an important source of the waverider lift in low-speed flight. • Aerodynamic performance of the cuspidal waverider in low speed range is the best. [ABSTRACT FROM AUTHOR]

Abstract The combined thermal protection system has led the drag and heat reduction of hypersonic reentry vehicles to a new developing direction. In order to obtain better resistance to aerodynamic drag and heat while maintain the work stability of the aircraft at the same time, the optimization design of the combined thermal protection system is indispensable. In this paper, a CFD numerical simulation method is combined with a surrogate model based multi-objective optimization algorithm to optimize the configuration of a combinatorial spike and opposing jet thermal protection system in a hypersonic flow with the freestream Mach number of 5.75. The obtained results show that when the total drag coefficient (Cd) and the heat fluxes on the head of the blunt body (Q) are the objective functions, the diameter ratio of the aerodisk to the blunt body bottom (d/D) has no significant effect on the drag and heat flux reduction, and it can be neglected in the optimization process. While the objective functions Cd and Q are affected by the three variables the length-to-diameter ratio of the aerospike (L / D), the jet pressure ratio (PR), and the nozzle radius (r 0) in similar ways. Therefore, in this case, the optimization problem can be transformed into a single-objective optimization problem. The quadratic response surface model established by sample points obtained by the orthogonal experimental design method is of high simulation accuracy, with the determination coefficients of Cd and Q are 0.964 and 0.965 respectively, and there is only a difference of −6.96% in the objective function Cd , −0.93% in Q between the optimization results and the CFD results. The combined thermal protection system has better performance in both drag and heat reduction than the single spike or opposing jet systems. Compared with the pure blunt body, the total drag coefficient and heat flux on the head of the blunt body with the combined thermal protection system have significantly decreased, with the Cd goes down 86.66%, and the wall heat flux Q drops off 96.37%. Highlights • The combinational concept was optimized by surrogate-based optimization approaches. • The combinational concept has better performance in both drag and heat reduction. • The optimized results match well with the numerical results. • All three design parameters have significant impacts on the drag and heat flux reduction. [ABSTRACT FROM AUTHOR]

Abstract According to the geometrical relationships of osculating cone waveriders, two kinds of design methods for the constant swept waverider are discussed in the current study, and both of them own more flexible design curves. The correctness and effectiveness of the proposed approaches has been verified by means of the CFD approach. The design Mach numbers of the cuspidal waverider and the general osculating cone waverider with the same volumetric efficiency of the delta-winged waverider have been determined by writing a program for calculating the volumetric efficiencies of waveriders. Furthermore, the high speed aerodynamic performance advantages of the cuspidal and delta-winged waveriders have been analyzed in detail, in comparison with the high speed aerodynamic performance of the general osculating cone waverider. The obtained results show that compared with the general osculating cone waverider, the cuspidal waverider has better high speed aerodynamic performance under any flight condition in the high speed range, while the high speed aerodynamic performance of the delta-winged waverider is worse. In addition, at the high flight Mach number, the cuspidal waverider shows the significant non-linear behavior of the vortex generated lift, while the delta-winged waverider and the general osculating cone waverider do not have this property. Highlights • Two kinds of design methods for the constant swept waveriders were proposed. • The two waveriders were generated based on the osculating cone approaches. • High speed aerodynamic performances of the two waveriders were analyzed and compared. • Cuspidal waverider owns better aerodynamic performances in high speed range. • Non-linear behavior of the vortex generated lift is shown in the cuspidal waverider. [ABSTRACT FROM AUTHOR]

In the current study, a variable Mach number waverider design approach has been proposed based on the osculating cone theory. The design Mach number of the osculating cone constant Mach number waverider with the same volumetric efficiency of the osculating cone variable Mach number waverider has been determined by writing a program for calculating the volumetric efficiencies of waveriders. The CFD approach has been utilized to verify the effectiveness of the proposed approach. At the same time, through the comparative analysis of the aerodynamic performance, the performance advantage of the osculating cone variable Mach number waverider is studied. The obtained results show that the osculating cone variable Mach number waverider owns higher lift-to-drag ratio throughout the flight profile when compared with the osculating cone constant Mach number waverider, and it has superior low-speed aerodynamic performance while maintaining nearly the same high-speed aerodynamic performance. [ABSTRACT FROM AUTHOR]

The major challenge among a number of design requirements for hypersonic vehicles is the reduction of drag and aerodynamic heating. Of all these techniques of drag and heat flux reduction, application of forward facing aerospike conceived in 1950s is an effective and simpler technique to reduce the drag as well as the heat transfer rate for blunt nosed bodies at hypersonic Mach numbers. In this paper, the flow fields around a blunt cone with and without aerodisk flying at hypersonic Mach numbers are computed numerically, and the numerical simulations are conducted by specifying the freestream velocity, static pressure and static temperatures at the inlet of the computational domain with a three-dimensional, steady, Reynolds-averaged Navier-Stokes equation. An aerodisk is attached to the tip of the rod to reduce the drag and heat flux further. The influences of the length of rod and the diameter of aerodisk on the drag and heat flux reduction mechanism are analyzed comprehensively, and eight configurations are taken into consideration in the current study. The obtained results show that for all aerodisks, the reduction in drag of the blunt body is proportional to the extent of the recirculation dead air region. For long rods, the aerodisk is found not that beneficial in reducing the drag, and an aerodisk is more effective than an aerospike. The spike produces a region of recirculation separated flow that shields the blunt-nosed body from the incoming flow, and the recirculation region is formed around the root of the spike up to the reattachment point of the flow at the shoulder of the blunt body. The dynamic pressure in the recirculation area is highly reduced and thus leads to the decrease in drag and heat load on the surface of the blunt body. Because of the reattachment of the shear layer on the shoulder of the blunt body, the pressure near that point becomes large. [ABSTRACT FROM AUTHOR]

Sufficient mixing between the supersonic airstream and the injectant is critical for the design of scramjet engines. The information in the two-dimensional supersonic jet-to-crossflow flow field has been explored numerically and theoretically, and the numerical approach has been validated against the available experimental data in the open literature. The obtained results show that the extreme difference analysis approach can obtain deeper information than the variance analysis method, and the optimal strategy can be generated by the extreme difference analysis approach. The jet-to-crossflow pressure ratio is the most important influencing factor for the supersonic jet-to-crossflow flow field, following is the injection angle, and all the design variables have no remarkable impact on the separation length and the height of Mach disk in the range considered in the current study. [ABSTRACT FROM AUTHOR]