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全乘波背部进气高超声速飞机新概念构型研究
Alternative TitleStudy of New Conceptual Configuration for Waverider-Based Hypersonic Airplane with Dorsal Intake Layout
肖尧
Thesis Advisor余西龙 ; 崔凯
2018-05-23
Degree Grantor中国科学院大学
Place of Conferral北京
Subtype博士
Degree Discipline流体力学
Keyword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
Abstract

吸气式高超声速飞行器是目前国内外的热点研究对象。随着相关研究技术(如超燃冲压发动机、热防护材料、飞行控制系统等)成熟度的不断提高,各类具有实用化前景的飞行器必然成为下一步的研究重点,其中具有远程快速运载功能的巡航型吸气式高超声速飞机就是主要方向之一。依据进气方式,高超声速飞机可主要分为腹部进气、旁侧进气和背部进气三种布局,其中,背部进气布局方案的特点是将进排气系统置于飞行器上部。由于推进系统与飞行器的主升力面解耦,可有效改善整机的气动特性,同时拓宽优化设计空间,因此近年来已成为一个重要的高超声速气动构型研究方向。

基于上述背景,本文以背部进气高超声速飞行器布局为主要立足点,以提升整机宽速域气动性能为主要目标,提出了一类多乘波体组合前体及整机概念,并从参数化设计方法、气动性能优化、宽速域特性分析等几个主要方面开展了一系列探索性研究。旨在为长航时高超声速飞行器构型设计提供一条新思路,并为下一代飞行器的研究提供有价值的设计参考。本文主要工作简介如下:

1)由于本文主要采用计算流体力学(CFD)手段对构型进行气动性能分析,为验证CFD计算的可靠性,以典型空天飞机标模和风洞实验数据为基准,分别采用结构网格与棱柱网格两种不同的拓扑形式,在马赫数0.5~8的宽速域范围内均开展了CFD可靠性验证计算。结果表明,在高超声速条件下,采用棱柱网格阻力系数与结构网格计算所获得阻力系数偏差较小,且CFD结果与实验数据的拟合度较高;在亚跨超条件下,结构网格计算结果具有较好的精度。

2)基于前体的设计需求,发展了一种将乘波体分为内/外两侧并分别采用前/后缘线对其定义的组合前缘乘波体设计方法。在乘波体中心区域定义直线前缘线,正流线追踪获得乘波面;在乘波体两侧区域定义直线尾缘线,逆流线追踪获得乘波面;两个区域相交部分共用同一条流线。采用均匀实验设计方法,在给定的设计空间内生成两组构型实例。结合CFD分析与构建代理模型的手段可得,运用该方法生成的乘波体可在较大参数范围内获得较好的升阻比,且外形相对平坦的乘波体构型其升阻比较优。

3)为满足背置双发布局的设计需求,发展了一种三乘波体前体的生成方式。CFD分析结果表明,此类前体构型于下侧均可保持很好的乘波特性,上侧由于上表面产生横向流动从而破坏其乘波特性,且随着拼接角度增大干扰越明显。另外,构型于出口处均可保持较好的流场均匀性,随着拼接角度增大其捕获流量越大;但随着拼接角度增大,其出口流场受攻角变化的影响越大。在此基础上,进一步开展了初步的翼身融合整机设计。分析结果表明,在考虑粘性的情况下,构型下表面均保持较完整的乘波特性,从而有效提升整机气动性能。构型最优升阻比均可达4.8以上(包含底阻),在小攻角范围内均可获得较优的升阻比。不同翼展构型具有一致的升力系数。此外,这类构型均为纵向静稳定。

4)开展了背置单进气口布局的构型研究。这类构型的乘波压缩面通过定义前缘线生成,并分三段来定义。机身同样采用翼身融合的设计。CFD计算结果表明,该类构型同样可获得较高的升阻比,且仍在小攻角工况范围内气动性能较优。前体对来流质量流量具有较好的捕获能力,于设计工况基本实现激波封口,出口马赫数分布均匀,受攻角、侧滑角的影响较小。另外,此类构型较易实现纵向和航向的静稳定性设计,利用乘波体的自然外凸亦可实现横向静稳定。

5)以升阻比最大作为优化目标,基于背置单进气口布局构型,结合运用CFD分析、均匀实验设计方法、代理模型方法和遗传算法开展了气动外形优化设计。结果表明,经优化后0°飞行攻角条件下构型升阻比可达5.6,表明了优化设计的有效性。同时基于代理模型开展了参数灵敏度分析。此外,对最优升阻比构型进一步开展了多攻角、侧滑角工况的气动性能评估。结果表明,其最优升阻比于2°飞行攻角条件下可达6.55。

6)在前述研究基础上,结合一种典型的TBCC发动机模块,以未来高超声速客机为应用设想,参考已有客机的相关参数,开展了具有实用化前景的飞行器概念研究,并比较了几种不同压缩面形状的构型。结果表明,即便在考虑发动机内流道阻力的条件下,采用背置单进气口布局的构型仍可获得较高的升阻比。同时,通过压缩面形状变化,可有效改善内部容积,且升阻比仅有小幅降低。另外,压缩面形状变化还对整机压心位置有较为明显的影响,因此也可作为飞行器配平姿态调整的有效手段。

7)考虑飞行器蒙皮连接间外形微小误差,结合本文构型较为扁平的特点,基于二维模型开展了飞行器表面接缝热流特性分析。分析结果表明,接缝尺寸越大,局部平均热流的突变较小;接缝凸起将导致局部热流的增加,凹陷则使局部热流降低,热流变化值与凸起/凹陷厚度基本呈线性关系;随着边界层厚度的增加(即接缝位置后移),接缝局部热流值的变化下降;当两个接缝间距较小时,靠后的接缝局部热流值降低。

现有结果表明,基于这种概念,由于飞行器下压缩面可以采用完整的乘波体设计,因此其在高超声速巡航条件下升阻比可以有效提升,应具有较好的实用化前景。

Other Abstract

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.

 

Call NumberPhd2018-026
Language中文
Document Type学位论文
Identifierhttp://dspace.imech.ac.cn/handle/311007/73138
Collection高温气体动力学国家重点实验室
Affiliation中国科学院力学研究所
Recommended Citation
GB/T 7714
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.
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