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The problem of a typical supersonic mixed compression inlet that is
characterized by multiple oblique shocks reflected in the supersonic
diffusser and a terminating normal shock standing in the subsonic diffuser
downstream of the throat is difficult to study. The complexities in this
knid of intake system of a high-speed vehicle are numerically
investigated. The interaction between shocks and the boundary layer plays
an important role in the development of the boundary layer which in turn
influences the performance of the inlet. The repid increase in the
boundary layer thickness due to a large adverse pressure gradient across
the shocks may cause flow separation, which can be eliminated by a careful
design of the inlet contour and the incorporation of a bleed system. For
the inlet studied here, the static pressure ratio through the inlet
passage is as high as 31 for a designed flight Mach number 3.0.
An improved scheme based on the second-order implicit upwind scheme of
Coakley''s is presented. The scheme is a finite volume form, solving the
Navier-Stokes equations. Several test cases are carried out to show the
accuracy and efficiency of the present method. Calculations of an
inviscid, one-dimensional, transonic nozzle flow and of a two-dimensional
transoic channel flow show better shock-capturing capability of the scheme
than a conventional central differencing scheme, or a flux vector
splitting upwind scheme, and are comparable to the TVD scheme. Further
verifications of the scheme both on the static pressure distribution and
the skin friction distribution along the wall in the shock/laminar flat
plate boundary layer interaction problem reveal that the present scheme is
accurate and stable. The flowfields of an oblique shock impinging on a
convex surface for different types of flow (i.e.,laminar, transitional,
and trubulent) and different bleed amounts and locations are investigated
for a better understanding of the effects of bleeding for boundary layer
control.
The objective of simulating a realistic mixed-compression supersonic inlet
is finally accomplished. Problems encountered in the simulations, such as
inlet unstart and overspeed analysis, are preseted and discussed in
detail. The detailed flowfields, surface pressure distributions, the
effect f bleed, simulation of a vortex generator (by enhancing a large
eddy viscosity can prevent the flow reversal at engine face) and inlet
performace parameters such as the total pressure recovery and the flow
distortion are presented. The predictions are good in comparing the
results with experimental data. It is believed and recommended that the
present method, due to its accuracy and robustness, can be used as a good
tool for the design of a supersonic intake system.
對於一具有多重斜震波反射於超音速擴散器及正震波形成於喉部區後次音速擴散器內
的超音速進氣道流場的分析是一件困難的工作。本文利用數值方法，藉以研究此類進
氣道系統的複雜流場。由於震波與邊界層之間的交互作用對邊界層的成長構成主要的
影響因素，因此，也進而影響到進氣道的性能。當邊界層承受由震波所引起之較大的
反向壓力梯度時，邊界層極容易流離物體表面，而此種現象可以以改進的內部外形設
計及有效利用吸氣裝置予以消除。針對本文所計算的設計點馬赫數為3.0 的進氣道而
言，出口與入口的靜壓比可以高達31.5，因此可見送向壓力梯度的重要性。
本文根據Clakley 所發展的二階準確隱式上游數值法，針對其缺點提出改良的方法，
並採用有限體積法之技巧以解可壓縮的Navier-Stokes 方程式。為了顯示本法的效率
及其準確性，測試相關的問題。對於非黏性一維及二維穿音速管流的問題，展現出本
法對於解析震波之能力，優於傳統之中央差分法及通向量分叉上游法，並且可以與先
進之TVD 法比擬。進一步的黏性流模擬於震波與平板層流邊界層交互作用之問題，更
顯現本法之準確性及穩定性。
最後本文完成了模擬一實際混合壓縮進氣道的流場及其分析。詳細的流場分析，壓力
在壁表面的分佈，吸流對於控制邊界層及正震波之穩定性的影響，渦流產生器的數學
模型模擬以及各種進氣道性能的參數指標，如全壓回復及流體扭曲都有詳盡的研究。
結果與實驗值所量測比較都相當吻合。我們相信並推薦本法，由於它本身的準確性及
穩定性，實可作為研究超音速進氣道系統之一有效分析工具。
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