宽速域来流对超声通流风扇叶型气动性能的影响

doi:10.7527/S1000-6893.2023.28523

Abstract

Abstract:

It is of great significance to understand the influence of wide speed inflow on the aerodynamic performance of Supersonic Through-Flow Fan （STFF） cascades to ensure the efficient and stable operation of the STFF. In this paper， the influence of inlet Mach numbers ranging from 0.10 to 2.36 on the aerodynamic performance of the STFF cascade was studied by numerical simulation. The flow structure evolutions of the STFF cascade with large negative incidences and critical Mach numbers were emphatically discussed. The results are as follows. The change from subsonic to transonic incoming flow caused a passage shock in the flow passage and increased the shock loss. Simultaneously， the interaction between the shock and the boundary layer of the adjacent blade suction surface induced large-scale flow separation and thus increased the viscosity loss. The evolution of the flow field structure with increasing negative incidences from -2° to -4° at a Mach number of 0.6 is similar to the above process. The difference was that no flow separation occurred on the blade surface in the latter case. The change of incoming flow from transonic flow regime under the condition of “unique incidence” to supersonic flow regime under the condition of 0° incidence slightly enhanced the shock intensity of the suction side branch of the leading-edge shock. Meanwhile， the pressure side branch of the leading-edge shock moved upstream resulting in a shock angle increase， while the Mach number ahead of the shock decreased. These two factors led to an intensity reduction of the pressure side branch of the leading-edge shock. Therefore， under the influence on both the suction and the pressure side of the leading-edge shock， the shock loss was eventually reduced. The interference between the pressure side branch of the leading-edge shock and the boundary layer of the adjacent blade suction surface was weakened， reducing the viscosity loss. In the cases of Ma=1.96 and 2.16 with an incidence varying from the negative critical value to the negative stall value， the strong interference between the branch on the suction side of the leading-edge shock and the pressure surface of the adjacent blade induced a flow separation. Meanwhile， the “virtual pressure surface” was established by the separation and the core flows. As a result， the suction side branch of the leading-edge shock changed from regular reflection to Mach reflection， and the suction side branch of the leading-edge shock near the pressure surface evolved into a Mach stem， increasing the shock loss.

Key words: axial supersonic through-flow fan, wide speed range, Mach number, loss characteristics, shock structure, flow separation

CLC Number:

V231

Shijun SUN, Xiaolong LI, Yanming LIU, Jianhua WANG, Songtao WANG. Influence of wide-speed-range inflow on aerodynamic performance of supersonic through-flow fan cascade[J]. Acta Aeronautica et Astronautica Sinica, 2023, 44(21): 528523-528523.

Figures/Tables 18

Fig.1

Table 1

Fig.2

Fig.3

Fig.4

Fig.5

Table 2

Inlet/outlet boundary conditions

来流马赫数

进口边界条件

出口边界条件

亚声速

（ $M a < 0.65$ ）

总温、总压与进口气流角

亚声速

跨声速

（ $0.65 ≤ M a < 1.58$ ）

总温、总压与进口气流速度分量

亚/超声速

超声速

（ $M a ≥ 1.58, M a z ≥ 1$ ）

总温、总压与进口气流角

超声速

Table 2

Fig.6

Fig.7

Table 3

Fig.8

Fig.9

Fig. 10

Table 4

Table 5

Fig.11

Fig.12

Fig.13

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参数	取值
安装角/（º）	22.6
进口金属角/（º）	32.6
出口金属角/（º）	11
设计攻角/（º）	0
来流湍流度/%	1

来流马赫数	可用攻角范围/（º）	最小损失
1.76	14.30	0.086 5
1.96	20.50	0.092 1
2.16	22.25	0.100 5
2.36	22.50	0.113 3

来流马赫数	攻角变化范围/（°）	总压损失变化率Δω/%
1.96	-10~-12	41.0
2.16	-12~-14	35.6

来流马赫数	总压损失系数
0.64	0.062 8
0.65	0.275 6
1.57	0.273 8
1.58	0.087 5

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Influence of wide-speed-range inflow on aerodynamic performance of supersonic through-flow fan cascade

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