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Free Stream xiao xiao. Free Stream Xiao q r. Free Stream xiao ting. Solo vine por el DDO y tambien por que me gusta la musica. Saludos desde Bolivia. Free Stream xiao hui. We use cookies to offer you a better experience, personalize content, tailor advertising, provide social media features, and better understand the use of our services. To learn more or modify/prevent the use of cookies, see our Cookie Policy and Privacy Policy. A design methodology has been developed for the base flow of streamline-traced hypersonic inlet. Limited by the inlet length, the truncated Busemann flow results in a curved shock wave accompanied with large total pressure loss near the axis. ICFA is then utilized to maintain the shock shape as straight as possible, downstream which an internal flow defined by spline surface is optimized for maximum total pressure recovery. Though the shock wave still curves, it only occurs in a very small region close to the axis. Accordingly, compared to truncated Busemann flow with same length and contraction ratio, the developed base flow contributes to a higher total pressure. Nomenclature a = the speed of sound r = radial distance from the axis Vr = radical velocity Vω = tangential velocity γ = specific heat ratio ω = angle to the axis American Institute of Aeronautics and Astronautics 1 Design of Base Flow for S tream line-T raced Hypersonic Inlet Lianjie Yue 1, Yabin Xiao 2, Lihong Chen 3, and Xinyu Chang 4 Institute of mechanics, Chinese Academy of Sciences, Beijing China, 100190 A design methodology has been developed for the base flow of streamline-traced hypersonic inlet. Nomenclature a = the speed of sound r = radial distance from the axis V r = radical velocity V ω = tangential velocity γ = specific heat ratio ω = angle to the axis I. Introduction or best hypersonic airbreathing engine performance, the inlet is a critical comp onent to provide a large amount of efficient compression. Design concepts for high performance scramjet inlets were well documented in the literature, which can be characterized in terms of the planar inlets and the inward-turning inlets. In planar inlets, the flow compressions are achieved by a series of planar or quasi-planar shock waves (ramp com pression inlet [1] sidewall compression inlet [2] and 3D inlet [3. Whereas, the inward turning inlets are designed to capture a portion of a known flowfield by specifying a leading edge and tracing the capture-perimeter stream lines through the baseline flowfield to create the inviscid waverider surfaces [4, 5. Designed for full mass capture, the inward turning inlets avoid additional drag due to the spillage at design condition. The Busemann inlet has received considerable attention lately as a classical streamline-traced airbreathing inlet [6, 7. Whereas it has very high inviscid pressure recovery, even at off-design operation [8] it is longer than the traditional outward turning inlet with the same capture area. Lim ited by the r estriction of overall weigh t and size of the scramjet, it is always needed for an inlet designer to compromise between the size and the performance of an inlet. Therefore in the practical application, the inlet was generally shortened by truncating the isentropic compression surface [9-11. Nevertheless, the flow pattern would be altered after truncation, especially near the axis. The coalescence of conical compression waves at the origin can not be realized and there is a curved shock wave that intersects with the axis at a point downstream the origin. The curvature of t he shock wave would result in a higher total pressure loss near the axis, even a normal Mach disk. The flow then becomes non-uniform at the ex it [12. Undoubtedly, the base flowfield has dominant effect on the inlet performance. The performance of truncated Busemann inlet would be degraded. The aim of this paper is to develop a better base f low to improve inlet performance. 1 Associate Professor, Key Laboratory of High Temperature Gas Dynamics, yuelj@im 2 Assistant Professor, Key Laboratory of High Temperature Gas Dynamics. 3 Associate Professor, Key Laboratory of High Temperature Gas Dynamics, Member AIAA. 4 Professor, Key Laboratory of High Temperature Gas Dynamics, Member AIAA. F 16th AIAA/DLR/DGLR International Space Planes and Hypersonic Systems and Technologies Conf AIAA 2009-7422 Copyright 2009 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. American Institute of Aeronautics and Astronautics 2 II. Truncated Busemann Flow The Busemann inlet should be shortened without severe deterioration in performance. Previous researches have analyzed the effect of the truncation angle on the base flowfield in order to determine the suitable truncated location. For incoming flow Mach number of 5. 05 and exit Mach number of 3. 0, a 5 truncated inlet provides a better performance. Its length is about 2/3 that of the full Busemann inlet, whereas the inviscid total pressure recovery has been maintained up to 90. 02. Further increase in the angl e would contribute to a rapid increase in total pressure loss, while only a little decreases in the compression length [12. Fig. 1 shows the flow patterns for the 5 truncated Busemann flow. By comparison, Fig. 2 illustrates the schematic of the full Busemann flow field. Without truncation, the axisymmetric internal flow consists of a set of conical Mach waves followed by a free-standing conical shock wave. All the conical Mach waves and the single freestanding conical shock wave coalesce at the apex “O”. The inlet takes the uniform parallel flow since the freestanding conical shock wave is canceled at the body shoulder. After truncation, the sharp leading edge corner results in the formation of an oblique shock wave. Moreover, the compression wa ves from the curved surface would not coalesce at the origin “0”, but intersect with the shock wave. A curved shock wave occurs that intersects with the axis at a point downstream the origin ‘0. The reflected conical shock wave also im pinges at a location downstream the shoulder rather than the shoulder in the non-truncated case. Although the total pressure recovery is still high, the curvature of the shock wave results in the high total pressure loss near the ax is, even a normal Mach disk. As shown in the Fig. 1, the total pressure recovery near the axis has come down to 65. 7% rather than 97. 2% that is far from the axis. This phenomenon decreases the total pressure; on the other hand, it degrades the flow uniformity. Correspondingly, if we can control the compression waves to maintain the shock strength, above di sadvantages can be perhaps limited. X Y -2 -1 012 0 0. 5 1 ptr: 0. 68 0. 72 0. 77 0. 82 0. 86 0. 91 0. 95 1. 00 a) Total pressure contour Y -2 -1 012 0 0. 5 1 ma: 3. 00 3. 29 3. 57 3. 86 4. 14 4. 43 4. 71 5. 00 b) Mach number contour Figure 1. Flow field of 5 truncated Busemann flow (incoming flow Mach 5. 05) American Institute of Aeronautics and Astronautics 3 Figure 2. Schematic of full Busemann flow III. Internal Conical Flow In order to control the shock strength along radical direction, we had better maintain the shock wave as straight as possible to avoid the larger shock angle around the axis. S. Molder has ever demonstrated four types of axisymmetric conical flow [13. Except the external conical flow, Busemann flow, ICFA (Internal Conical Flow A) was analyzed as a convergent f low, in which parallel flow passes through a conical shock and converges towards the axis, as shown in Fig. 3. This flow can also be described by the Taylor-Maccoll equation. 2 2 2 2 2 / 1 2 1 a V V V ctg V V V r r r r r − ? ? ? ? ? ? ′ − − ? ? ? ? ? ? ′ + − = ″ ω ω γ (1) Where V r is the velocity in the r direction, nondimensionalized with respect to the escape speed. Figure 3. Schematic of internal conical flow A (ICFA) The ICFA flow starts from a straight internal conical shock which leads to constant total pressure recovery behind the shock, which will contribute to the flow uniformity of the total pressure recovery within the inlet. Unfortunately, from the equation, we see that there is so-called singularity condition wh en the denominator equals 0. The numerical solutions shows that near the singularity V ω approaches the speed of sound, and consequen tly V r approaches infinity. It was noted also that near the singularity the surface starts turning extremely sharply further away from the freestream flow. The existence of the singularity indicates that ICFA can not be realized completely and regular shock reflection at the axis is impossible. Nevertheless, we can expect to maintain the initial shock wave as straight as possible by use of ICFA. The singular ray is a characteristic line because the tangential velocity is the speed of sound. Therefore the shape of initial shock would be only influenced by the surface upstream the singularity. In orde r to obtain a nearly straight shock wave, the body surface of ICFA can be specified till a location “A” close to singularity point “S”. A curve is subsequently connected from point “A” to t he throat, as shown in Fig. 4. Though the shock pattern would be interfered by the surface downstream the connecting point “A”, the eff ect would be limited within a small region close to the axis. American Institute of Aeronautics and Astronautics 4 IV. Optimization Method and Numerical Setup The connecting curve AD is also important for the base flow performance, which will influence the initial shock around the axis and the reflected shock wave. In this paper, optimization was utilized to search the best curve with maximum total pressure recovery. In order to compare with the truncated Busemann flow, the length and contraction ratio of optimized base flow were specified to equal to those of the truncated Busemann flow. Meanwhile, the angle of shock emanating from
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