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Tip leakage flow is an important factor affecting the efficiency and stability of compressors. It is important to study the tip leakage flow model in the design and performance prediction process of turbomachinery. This work analyzes the existing tip clearance loss prediction model and refines it to predict the tip clearance loss considering more blade geometric parameters. After analyzing the Yaras tip leakage loss model derivation process, it is found that the discharge coefficient (taken as constants by Yaras) will change with the gap geometry, namely, it is related to the maximum thickness and the gap size of the blade. In this paper the discharge coefficient of tip gap is revised to better predict the effect of gap loss with regard to the maximum thickness of the blade. In this paper, the research objects are linear cascades with NACA65 profile and numerical experiments were carried out by a CFD package Numeca under the condition of Mach number equal to 0.45, where the variables are the gap sizes (1%, 2%, 3%, 4%, 6% of the axial chord length) and the maximum thickness of blades (3%, 5%, 7%, 9%, 11% of the axial chord length). Combined with the numerical calculation results, according to Yaras loss prediction formula and using the similar characteristics to the discharge coefficient of the variation trend and the Planck blackbody radiation formula, the relationship between the discharge coefficient and the maximum thickness of the blade and the tip clearance is summarized, which is integrated with the Yaras loss prediction formula to obtain the final formula. This prediction formula fits the NACA65 blade calculation fairly. The average error of 24 calculation points is 2.59%. Then the improved model was compared with several existing models. Besides the CDA062, C4 and polynomial thickness distribution blade tip clearance loss is predicted using the refined prediction formula and the biggest prediction error is 5.46%. Therefore, the improved formula still has a good prediction effect when the blade type change. It is considered that within a certain range of maximum thickness of the blade type and tip gap sizes, the improved formula can give good predictions even though the blade type is different.

At present, researches on the discharge coefficient of the combustion chamber cooling holes are mostly based on the experimental study of the porous plates considering various geometric structures under different flow conditions. In this method, the average discharge coefficients of multiple holes are obtained. Since the discharge coefficient will be applied to the numerical simulation, it is important to obtain an accurate formula for each cooling hole. Therefore, the discharge coefficient will be associated with some local aerodynamic parameters around the cooling holes and geometric parameters of the cooling holes. In this paper, the geometric parameters consist of length-to-diameter ratio (L/d) and inclination angle (α). The aerodynamic parameters cover the Mach number of cooling flow and mainstream (Mac, Mam), characterizing the flow feature, and the pressure ratio (π) which associates the cooling flow with the mainstream. The purpose of this paper is to study the discharge coefficient of a single circular hole with variable geometries under a cold condition of a combustion chamber, which has low crossflow and low-pressure ratio on both sides of the hole. In this environment, according to the research results, the discharge coefficient is sensitive to the cooling flow Mach number, length-to-diameter ratio and pressure ratio (π ≈ 1.05). Discharge coefficient decreases with L/d linearly, conforms the quadratic function with Mac and changes complexly at π ≈ 1.05. Other parameters have little effect on the discharge coefficient. The data for discharge coefficient of the cylindrical hole considering different parameters is obtained through numerical simulation and the correlations summarized by these data are valid for the following ranges: L/d = 3∼12, α = 20°∼45°, Mac = 0.05∼0.15, Mam = 0∼0.1, π = 1.05∼1.15. Compared with the CFD data, the prediction formula has a maximum error of less than 3% and a mean absolute error of 0.78%.

Through-flow methods are widely used in turbomachinery design and performance prediction due to its fast calculation. In the past, circumferentially-averaging through-flow methods tended to ignore circumferential fluctuation source term (CFST) in the governing equations, leading to an inaccurate result when blades showed a strong three-dimensional feature. To solve this problem, CFST needs to be properly modeled. In this paper, three-dimensional numerical simulation is conducted to find the area where CFST dominates. The influence of inlet Mach number, incidence, camber angles and sweep angle on the CFST in this area is analyzed. A model for CFST is constructed. Three-dimensional verification of this model suggests that within the common scope of use of aerospace compressor designing, this model shows a good accuracy.

Swirl inlet distortion is usually encountered in modern flight vehicles since their inlet ducts usually consist of one or two bends, such as S-inlet duct. An experimental device is first designed to simulate the swirl inlet distortion and then used to test the effectiveness of a novel casing treatment (CT) on a low-speed compressor under the swirl distortions of various intensities. The influences of co- and counter-rotating swirl inlet distortion on the test compressor and the stabilization ability of this novel CT are well demonstrated by the illustrations of static pressure rise curves and efficiency curves. The dynamic prestall pressure signals are also captured to reflect the perturbation energy in the flow field through which the mechanism of the novel CT will be indicated. The relevant results show that counter-rotating swirl distortion in small intensity could increase the compressive ability of compressor with small efficiency loss, and the co-rotating swirl distortion always brings about detrimental effects on compressor performance. At the same time, the distortion of twin swirls can cause nonuniform total pressure profile which can seriously damage the compressor performance. Besides, the stall precursor-suppressed (SPS) CT shows a good capability of stall margin (SM) enhancement no matter what swirl inlet distortions are encountered in the test compressor.

This paper conducts an experimental research of rotating inlet distortion on a low-speed large size test compressor with emphasis on the stability problem of axial fan/compressors, and the stall margin enhancement with a kind of stall precursor-suppressed (SPS) casing treatment. Some results on compressor stall margin and prestall behavior under the restriction of rotating inlet distortion are presented. The experimental results show that whether the inlet distortion is co-rotating or counter-rotating, the SPS casing treatment can still improve the stall margin without leading to additional efficiency loss caused by such configuration. The experiment results also show that the mechanism of the stall margin improvement with such casing treatment is associated with delaying the nonlinear development of the stall precursor waves and weakening the unsteady flow disturbances in a compression system.