|英文题名||Study of New Conceptual Configuration for Waverider-Based Hypersonic Airplane with Dorsal Intake Layout|
|导师||余西龙 ; 崔凯|
|关键词||高超声速飞行器 乘波体 背部进气布局 升阻比 气动设计与优化|
Hypersonic Air-breathing Vehicle (HAV) research is an international hot issue nowadays. With the development of researches in scramjet, thermal protection materials, flight control system, etc., various kinds of HAV, which can be used in practical situations, will become the research priorities in the near future. One of the main research directions is the hypersonic air-breathing cruise airplane, which has the fast long-range carrying capability. According to the intake mode, the hypersonic air-breathing airplane can be mainly divided into three types of layouts: ventral intake layout, flanking intake layout and dorsal intake layout. The dorsal intake layout is characterized by placing the propulsion system on the upper of the airplane. Because the propulsion system is decoupled from the main lifting surface of the airframe, it can improve the aircraft aerodynamic performance effectively and broaden the design and optimization space at the same time. Thus the dorsal intake layout has become an important research direction for hypersonic aerodynamic configuration in recent years.
Based on the above background, focused on the dorsal intake layout of hypersonic airplane, taking the improvement of aerodynamic performance in the wide speed range as the main goal, a class of multi-waverider forebody and aircraft concept is proposed in this paper. In addition, a series of exploratory studies have been carried out from several aspects such as parametric design methods, aerodynamic performance optimization, and wide speed range aerodynamic characteristic analysis. The studies are aim at providing a new idea for the configuration design of long-endurance hypersonic vehicles and providing a valuable design reference for the research of next generation aircraft. The studies have been finished as follows.
This paper mainly uses Computational Fluid Dynamics (CFD) to analyze the aerodynamic performance of the aircraft. In order to verify the reliability of CFD simulation, taking the typical space shuttle model and wind tunnel experiment data as the benchmark, CFD reliability verification examples have been carried out in a wide speed range of Mach 0.5~8. Structured grid and the prism grid are used in those examples respectively. The results show that, under the hypersonic condition, the deviations of the drag coefficients that results from structured grid and prism grid are small， and the fitting degree between the CFD results and the experiment data is high. Under subsonic, transonic, and supersonic conditions, all of the structured grid results have good accuracy.
To meet the design requirements of the forebody, a multi-segment-leading-edge waverider generating method, which dividing waverider into inner/outer parts and defining leading/trailing edges respectively, is developed. A straight-line leading/trailing edge is defined in the inner/outer part of the waverider, and streamline tracing is used to obtain the waverider surface. The intersections of the two parts share the same streamline. Using uniform experimental design methods, two sets of configuration samples are generated within a given design space. Its aerodynamic results can be obtained by combining the CFD analysis and surrogate model building. The results show that waveriders generated by this method can obtain a well lift-to-drag ratio in a large parameter range, and the relatively flat configuration has better lift-to-drag ratio.
In order to meet the design requirement of the dorsal dual engine layout, a trinal-waveriders forebody generation method is developed. The results of CFD analysis show that the lower side of this type of forebody can well maintain the waverider characteristics; however, waverider characteristics is destroyed on the upper side of the forebody due to the lateral flow which is generated by the upper surface. As the assembled angle increases, the waverider characteristics destruction becomes more obvious. In addition, all of the forebodies can maintain good flow field uniformity at the outlet cross-section. Therefore, the mass flux rate captured by the forebody would be larger as the assembled angle increases. However, as the assembled angle increases, the impact on the outlet flow field, which caused by the change of angle of attack, is greater. On this basis, the preliminary configuration design of the blended wing body airframe is carried out. The analysis results show that the lower surface of the whole aircraft can completely maintain the waverider characteristics under the viscous condition. It improves the aerodynamic performance of the whole aircraft effectively. The maximal lift-to-drag ratio of the aircraft can reach 4.8 or more (including the base drag), and relatively high lift-to-drag ratio can be obtained within small angle of attack range. Furthermore, aircrafts with different aspect ratio have consistent lift coefficient. In addition, all of the aircrafts are statically stable in the pitching direction.
A configuration study for the aircraft with dorsal single inlet layout is carried out. The waverider of this type of aircraft is generated by defining the leading edge, which is defined in three sections. The airframe also uses the blended wing body design. The CFD analysis results show that this type of aircraft can also achieve high lift-to-drag ratio values, and it still obtains well aerodynamic performance within small angle of attack range. In addition, the forebody has a good ability to capture the mass flux, and the shock wave is attached to the cowl edge under design condition. The Mach number distribution is uniform at the outlet cross-section, and it is less affected by angle of attack and angle of sideslip. Besides, this type of aircraft is easy to achieve the statically stable design in the pitching and yawing direction, and the rolling statically stable design can be achieved by using the natural convex of the waverider.
Based on the aircraft with dorsal single inlet layout, a configuration optimization is carried out with the aim of maximum lift to drag ratio by combining with the uniform experimental design method, the computational fluid dynamics, the radial basis function surrogate model method and the genetic algorithm. The results show that the lift-to-drag ratio of the optimized aircraft can reach 5.6 under 0° angle of attack, which indicates the effectiveness of the optimization design. Moreover, the sensitivity analysis of the optimization results for each parameter is implemented. In addition, the aerodynamic performance evaluation for the optimized aircraft is carried out under different flight conditions such as different angle of attack, angle of sideslip.
Based on the former studies, combined with a typical Turbine Based Combined Cycle (TBCC) engine module, taking the future hypersonic airliner as an application envisagement, referring to the relevant parameters of the existing airliner, a study on the concept of the aircraft, which has a practical prospect, is carried out. Three airplanes with different compression surface deformations are designed for aerodynamic performance comparison. The results show that this kind of airplane also can achieve high lift-to-drag ratio even with the consideration of the engine drag. Meanwhile, the internal volume can be effectively increased by using the compression surface deformation technology, and the airplane lift-to-drag ratio is only slightly reduced. In addition, the compression surface deformation also has a significant effect on the position of the pressure center of the airplane, therefore it can be used as an effective means to adjust the airplane trim attitude.
Considering the tiny shape error of the aircraft skin connection, combining the relatively flat features of aircrafts in this paper, the heat flow characteristics of the aircraft surface seams are analyzed based on the two-dimensional model. The analysis results show that the lager the seam size is, the smaller the alteration of the local average heat flux is. Seam protruding will lead to increased local heat flux, and seam concave will decrease local heat flux. Heat flux alteration is approximately linear with the thickness of the protruding/concave. With the increase of the thickness of the boundary layer (scilicet the position of the seam is moved backwards), the alteration of the local heat flux at the seam decreases. Besides, when the gap between two seams is small, the local heat flux value at the rear seam will decrease.
The above results show that because of the lifting surface of the aircraft can be designed as a complete waverider, the lift-to-drag ratio of this kind of aircraft can be improved effectively under hypersonic cruise conditions. Therefore, this aircraft design concept should have a good prospect of actual application.
|肖尧. 全乘波背部进气高超声速飞机新概念构型研究[D]. 北京. 中国科学院大学,2018.|