Transonic Axial Flow Compressor Research Articles

An Improved Body Force Model to Simulate Blade Loading Variation Under Various Operating Conditions in Transonic Axial Flow Compressor

Abstract In order to analyze the unsteady aerodynamic response of compressors with limited computational resources, the development of the body force (BF) model has been pursued for decades. This model can simulate the effects of the compressor blade on the flow by applying a three-dimensional (3D) force field. Indeed, the accuracy of the BF model simulation highly depends on the construction strategy of BF field. However, for a transonic compressor, the influence of shock waves on the spatial distribution of blade loading is significant, which makes the construction of an appropriate BF field challenging. To solve this issue, an accurate BF model is proposed to describe the effects of the blade on the flow with high spatial fidelity in transonic compressors. For verification, a typical transonic fan rotor, NASA-Rotor 67, is selected. The numerical results calculated by the new BF model are compared with those obtained by the Reynolds-averaged Navier–Stokes (RANS) solver. According to the results, the new model can accurately capture the distribution of blade loading in the transonic blade channel at various operating conditions, eventually leading to a good prediction of compressor performance and the radial distribution of flow parameters downstream of the rotor. Moreover, this study discusses the causes of modeling errors and presents the application of the model in inlet distortion simulation at the end.

Three-dimensional optimization of squealer-tip for a transonic axial-flow compressor rotor blade

This research represents an innovative method for geometry generation of the squealer-tip in axial compressor rotors and its exploit in a numerical optimization process to obtain a better stage performance. For this purpose, the NASA Rotor-67 transonic compressor rotor blade is used as a test case to study the aerodynamic performance using computational fluid dynamics. The validation was performed for the characteristic map at the design speed and the comparison with the experimental results indicates excellent matching and high adaptability of the numerical method. An ingenious method of producing squealer tip for an axial compressor rotary blade is presented in this article, which is capable for locally shaping both suction and pressure surface geometry at a desired spanwise location simultaneously, while keeping the tip clearance at its value of the baseline NASA Rotor-67 geometry. In this method, control points are used to produce the starting spanwise location of the squealer, and modify the depth of the squealer geometry. The L-27 orthogonal array of the Taguchi method as the Design of Experiment (DOE) has been used to investigate the sensitivity of the aerodynamic results in three performance points of the choke, design and near stall regions, in relation to the design variables of the squealer. The generated database in the sensitivity analysis was used to train artificial neural networks to replace the CFD solutions with overwhelming run time. By coupling the genetic algorithm to the aforementioned neural networks and by applying penalties to maintain the minimum performance of the Rotor-67, enhancement of total pressure ratio, adiabatic efficiency, mass flow rate and even the surge margin was achieved. The main effect of the squealer is to modify the shape of blade tip vortices, and by more dissipation of energy in blade tip area and reduced equivalent flow area in this region, finally results in improved overall mass flow rate, total pressure ratio, adiabatic efficiency and surge margin by 0.58 %, 0.36 %, 0.19 % and 4.81 % respectively, at design point.

Optimization of Three-Dimensional Blade and Variable Stators for Efficiency and Stability Enhancement of Multistage Axial Flow Compressor at Variable Speeds

Abstract High efficiency and wide stability at variable speeds are equally important to the design and operation of multistage axial flow compressors. However, published research works on three-dimensional design optimization of compressors are so far mainly limited to a single blade row or stage at design speed due to the curse of dimensionality. Moreover, optimization of variable inlet guide vanes (IGV)/stators for off-design operations is carried out by using a rapid but low-fidelity prediction tool and is generally independent of design optimization of blade geometry. To tackle these issues, a three-dimensional holistic design and adjustment optimization method is developed in which both three-dimensional blade geometry and variable IGV/stators are optimized simultaneously for better efficiency and stability at design and off-design conditions. Metamodel-interpreted data mining method and adaptive infilling strategy are used respectively to enhance the capability of the metamodeling and optimization. The developed method is then applied to a modern highly loaded 3.5-stage transonic axial flow compressor at both design and part-design speeds. The results show that the stall margin is extended from 8.23% to 19.65% at 70% design speed while peak efficiency is slightly improved at design speed. The flow mechanisms responsible for the efficiency enhancement at design speed are mainly associated with the reduced total pressure loss in stators as well as inter- and intra-stage loading redistribution. The stability enhancement at 70% design speed is mainly achieved by loading the front blade rows while unloading the limiting rear blade row through variable IGV/stators adjustment. The developed holistic design and adjustment optimization method with the aid of metamodel-interpreted data mining is of great application value for the design and adjustment of advanced multistage axial flow compressors.

Investigation of transitional flow in a transonic compressor rotor with hub leakage using large eddy simulation

The hub leakage flow has been acknowledged as an important factor for performance deficiency of axial-flow compressors. Meanwhile, the laminar-turbulent transition in compressors is highly sensitive to the upstream flow state and significantly affects the flow loss. In the present work, quasi-wall-resolved large eddy simulation of a transonic axial-flow compressor rotor at the near-peak-efficiency point is carried out to investigate the effects of hub leakage as well as its absolute tangential velocity on the compressor performance and the laminar-turbulent transition in the blade passage. It is confirmed that the hub leakage with an absolute tangential velocity of 0.5 wheel speed can result in the near-hub total pressure deficit. With the hub leakage taken into account, the predicted total pressure ratio and adiabatic efficiency agree well with the experimental data. The simulation results indicate that increasing the absolute tangential velocity of hub leakage would intensify the near-hub vortices, elevate the endwall turbulence level, increase the near-hub flow loss, and cause a remarkable total pressure ratio drop. This work promotes the understanding of complex flow mechanisms in axial-flow compressors in the presence of hub leakage.

Evolution of unsteady secondary flow structures near the onset of stall in a tip-critical transonic axial flow compressor stage

The aerodynamic stability of an axial compressor stage depends on the rotor and stator. However, due to the specific design requirements, the evolving flow field and the resulting secondary flow structures are different in both components. This study investigates the evolution of dominant secondary flow structures occurring in the rotor and stator of a tip-critical transonic compressor stage at the near-stall condition using unsteady numerical analysis and validates performance characteristics with the experimental data. The investigation reveals that the presence of rotor tip shock creates a large difference in the pressure gradient across the pressure surface and the suction surface, intensifying tip leakage flow and shock-induced boundary layer separation. The higher incidence angle near the hub leading edge creates a local separation and reattachment zone. The radial pressure gradient causes the low momentum flow from this local separation zone to migrate radially upward. This migrated flow interacts with the tip leakage flow and separated blade boundary layer, eventually creating a colossal recirculation zone and subsequent rotor blockage of around 46%. The increasing streamwise adverse pressure gradient pushes the tip shock upstream, and at the onset of the stall, the flow directly separates from the rotor leading edge avoiding the shock formation. The stator flow field is dominated by the asymmetric hub corner separation induced by the streamwise adverse pressure gradient and the tip corner separation caused by the vortex structures convected downstream from the rotor tip region leading to stator blockage of around 48%. Along with the blade passing frequency, four other dominating frequencies (0.06×BPF, 0.12×BPF, 0.44×BPF, and 0.84×BPF) related to the aerodynamic instabilities are observed at the inception of stall.

Sensitivity of Weighing Functions in Genetic Algorithm for Efficiency and Pressure Ratio Optimization in Transonic Axial Flow Compressor

Axial flow compressor design is a multi-objective problem with objectives such as maximization of efficiency, total pressure ratio, mass flow rate and minimization of weight. In this work, validated meanline design process is integrated with genetic algorithm to optimize the compressor design. Fitness function used to evaluate effectiveness of candidate design solutions is arithmetic summation of combination of objective functions and respective weighing functions. Objective functions considered are pressure ratio, efficiency and operating margin in terms of De-Haller Number and Diffusion Factor. The summation of all the weighing functions is unity in all the iterations carried out. Weighing functions of the objective functions used in fitness function dictate the performance. Sensitivity study of weighing functions for multi-objective fitness function has been conducted for the intended mass flow and total pressure ratio. Effect of sensitivity of weighing functions in multi-objective fitness function on efficiency and pressure ratio in transonic axial flow compressor design is studied and discussed. The study carried out shows maximum value of weighing function does not yield maximum value of the objective function. Design trial of weighing function combination of (0.3 pressure ratio and 0.3 efficiency) yields an optimum pressure ratio of 1.588 and 88.1% efficiency for optimized fitness function. The iterative data generated is useful for assigning weighing function value based on design requirement.

Cfd Analysis to Investigate the Effect of Leaned Rotor On the Performance of Transonic Axial Flow Compressor Stage

This paper describes the numerical study carried out to understand the impact of combination of forward and backward leaned rotor on the overall performance of a transonic axial flow research compressor of CSIR National Aerospace Laboratories (NAL), Bangalore. The analyses were carried out using commercial Computational Fluid Dynamics (CFD) solver Ansys-CFX. Initially CFD analysis was carried out for the baseline rotor configuration and validated the results with the experimental data. Then, three new configurations of combined Leaned rotors were modeled from the radially stacked baseline rotor by changing the circumferential curvature of original stacking line using three control points located at 0%, 50% and 100% of the blade span. Analyses were carried out for all the three configurations of modified geometries and the results were compared with the baseline configuration. The results revealed many interesting aspects which will be very useful for better understanding of the blade curvature effects on the shock and secondary losses within the transonic rotor. The discussions on the effect of combination of forward and backward leaned rotor and the consequent developments on the overall performance of the compressor are presented in detail in this paper.

Design optimization and flow analysis of discrete tip injection in a transonic compressor based on nonlinear harmonic method and endwall blockage attenuation

Discrete tip injection (DTI) shows great promise for improving the operating stability of transonic axial flow compressor (AFC) rotors. However, the design optimization of DTI remains a challenging task because of both the reliance on computationally expensive unsteady simulations to calculate its effects and the lack of a flow physics-based index for assessing operating stability. The present study introduces a nonlinear harmonic method for the rapid simulation of the dominant unsteady effects caused by a DTI device, and it proposes the unsteady shroud endwall blockage attenuation as an operating stability optimization index for DTI design based on analyzing the stall flow mechanism in transonic AFCs. On this basis, an efficient optimization method for DTI design is proposed in combination with an adaptive kriging-based optimization technique. This design optimization method is validated by the Coandă injector design for the transonic rotor National Aeronautics and Space Administration Rotor 37, with improved operating range and reduced injection mass flow pursued simultaneously by a comprehensive objective function. The optimal DTI design significantly reduces the stalling flow coefficient of the compressor by 4.46% at a small injection mass flow (0.72% of the compressor stalling mass flow), with a slight increase in the aerodynamic performance of the compressor. Detailed unsteady flow-field analysis shows that the main reason for the improved operating stability of the transonic AFC is a significant attenuation and delayed recovery of shroud endwall blockage, and the underlying flow mechanism is elucidated well.

Effect of Different Radial Inclined Angles of Self-circulation Casing Treatment on the Transonic Axial Flow Compressor Performance

Unsteady numerical simulation of single passage was carried out on NASA Rotor 35 to study the influence of the radial inclined angle of self-circulation casing treatment (SCT) on the performance and stability of a transonic compressor at 100% design speed. The radial inclined angles were set to 0°(D0), °(D30) and 60°(D60), respectively. The calculated result indicates an increase in the stall margin improvement (SMI) and the design efficiency improvement (DEI) as the radial inclined angle increases gradually. The SMI of the SCT with a 60° radial inclined angle is 12.5%, the biggest among the three SCTs, and the peak efficiency improvements (PEI) of the SCTs are almost the same. The radial inclined angle is provided with the effect of strengthening the self-circulation casing treatment effect, which further improves the stable working range of the compressor, and the efficiency loss is also lower than that of the 0° angle structure. The flow conditions inside the bleeding part can be improved by radially skewing the self-circulating structure toward the rotor rotating direction.

Vortex Structure Topology Analysis of the Transonic Rotor 37 Based on Large Eddy Simulation

Highly three–dimensional and complex flow structures are closely related to the aerodynamic losses occurring in the transonic axial–flow compressor. The large eddy simulation (LES) approach was adopted to study the aerodynamic performance of the NASA rotor 37 for the cases at the design, the near stall (NS), and the near choke (NC) flow rate. The internal flow vortex topology was analyzed by the Q–criterion method, the omega (Ω) vortex identification method, and the Liutex identification method. It was observed that the Q–criterion method was vulnerable to being influenced by the flow with high–shear deformation rate, especially near the end–wall regions. The Ω method was adopted to recognize the three–dimensional vortex structure with a higher precision than that of the Q–criterion method. Meanwhile, the Liutex vortex identification method showed a good performance in vortex identification, and the corresponding contribution of Liutex components in the vortex topology was analyzed. The results show that the high–vortex fields around the separation line and reattachment line had high vortex components in the x–axis, the tip clearance vortices presented a high–vortex component in the y–axis, and the suction side corner vortex possessed high–vortex components in the y– and z–axes.

Experimental investigation of instability inception on a transonic compressor under various radial distributions of loading

The radial distribution of blade loading could be various due to the inlet distortion, the change of rotor speed or mismatching between stages, which is a key factor affecting the flow stability. In this paper, the instability inceptions under the various radial distribution of loading are experimentally investigated on a transonic compressor. The research subject is chosen as a 1.5-stage transonic axial flow compressor in which three types of instability inceptions were previously detected, including spike-type inception initiated at the rotor tip under the low speed, partial surge initiated at the stator hub under the high speed, and hub instability when the slot casing treatment was mounted. In order to facilitate the conversion of the compressor instability inception, the rotating speed is selected (17,160 r/min in this study) at which the blade loading at both the stator hub and rotor tip reach the critical value near the rotating stall. Furthermore, the variation of radial distribution of loading is induced by the radial non-uniformity of total pressure. Meanwhile, the high-response dynamic test points are mounted to measure static pressures at rotor tips and total pressures at stator trailing edges during the experiments. To detect the system response, an additional high response sensor is installed on the plenum inner wall. As demonstrated by the WFT and filtering methods, the results indicate that the instability evolution varies gradually with the radial distribution of loading at the inlet. The partial surge instability converts into the spike with the increase of tip loading intensity. With the increase of hub loading intensity, the partial surge instability converts into the hub instability. For the hub instability, a special process is that a sudden decrease of total pressure occurs at the stator hub. Furthermore, the short-length scale disturbance occurs at the rotor tip, and leads to the final compressor stall.

Analysis of the Flow Field at the Tip of an Axial Flow Compressor during Rotating Stall Process Based on the POD Method

Rotating stalls are one of the most dangerous phenomena to be avoided in the designing and operating of axial flow compressors. An understanding of the evolution of the flow characteristics around the rotor tip region is important to study the process of stall development. In this paper, some critical characteristics of the stall-related structures, which could not be observed by the traditional analysis methods, are understood by using the proper orthogonal decomposition (POD) method. The detailed unsteady flow fields in a typical transonic axial flow compressor during the stall process are obtained by the validated numerical simulation. Thereafter, the proper orthogonal decomposition method, as a method for model reduction and decomposition, is adopted to extract the flow characteristics from the numerical results of the stall process. It is found that the flow characteristics during the stall process can be well decomposed by the POD method. The pre-stall POD results show that the important flow features can be extracted and revealed in the low-order modes, which are not obvious in the original flow field. When the stall cells are formed and developed, the flow characteristics are gradually determined by the modes, which are related to the features of the stall cells. When the compressor is operated under stable stall conditions, the low-order POD modes are composed of a series of harmonic modes, which are sinusoid-like in space and time with the frequency of the stall cell rotation.

An overall blockage attenuation-based aerodynamic performance and stability design optimization method for transonic axial flow compressors

PurposeThis study aims to use computational fluid dynamics (CFD) to understand and quantify the overall blockage within a transonic axial flow compressor (AFC), and to develop an efficient collaborative design optimization method for compressor aerodynamic performance and stability in conjunction with a surrogate-assisted optimization technique.Design/methodology/approachA quantification method for the overall blockage is developed to integrate the effect of regional blockages on compressor aerodynamic stability and performance. A well-defined overall blockage factor combined with efficiency drives the optimizer to seek the optimum blade designs with both high efficiency and wide-range stability. An adaptive Kriging-based optimization technique is adopted to efficiently search for Pareto front solutions. Steady and unsteady numerical simulations are used for the performance and flow field analysis of the datum and optimum designs.FindingsThe proposed method not only remarkably improves the compressor efficiency but also significantly enhances the compressor operating stability with fewer CFD calls. These achievements are mainly attributed to the improvement of specific flow behaviors oriented by the objectives, including the attenuation of the shock and weakening of the tip leakage flow/shock interaction intensity.Originality/valueCFD-based design optimization of AFC is inherently time-consuming, which becomes even trickier when optimizing aerodynamic stability since the stall margin relies on a complete simulation of the performance curve. The proposed method could be a good solution to the collaborative design optimization of aerodynamic performance and stability for transonic AFC.

Metamodel-Interpreted Data Mining for Stability and Efficiency Enhancement of Multistage Axial-Flow Compressors

Abstract The tradeoff between higher efficiency and wider stability of performance map is still one of the bottlenecks to hamper the further research and development of advanced multistage axial-flow compressor. The recent rapid growth of computational resources and artificial intelligence has enabled data mining as one of the most effective and potential ways to gain a deep insight into the complex correlations between aerodynamic performance and three-dimensional geometry parameters. In the open literatures, however, few research works have been found on using the data mining that is independent of design optimization to extract priori design guidelines for multistage axial-flow compressor mainly due to the lack of proper data mining method focused on the interpretation of metamodel with full use of limited time-consuming computational fluid dynamics dataset. To tackle this issue, a metamodel-interpreted data mining framework is developed in which extreme gradient boosting (XGBoost) metamodel combined with Shapley additive explanation (SHAP) model are employed to locally interpret the feature importance of each sample in the computational fluid dynamics dataset and then extract the design guidelines in terms of the most influential geometry parameters and their beneficial variation directions. The developed method is applied to data mining of design guidelines for efficiency and stability enhancement of a front 3.5-stage transonic axial-flow compressor in ship-board gas turbine usage. The results show that the aerodynamic performance of the investigated multistage compressor is most sensitive to three-dimensional geometry parameters related to blade lean, blade twist, and variable stators. Specially, the variable stators mainly affect the stall margin at part speed. The blade lean mainly influences the adiabatic efficiency at design speed as well as the stall margin at both speeds, while the blade twist mainly influences the aerodynamic performance at design speed. New designs followed by the design guidelines are obtained and critical performance indicators related to the goals of the data mining task are verified. The stall margin at part speed is widened to 5.87% with adjustment of blade lean and twist and further to 23.31% with additional adjustment of variable stators. The peak adiabatic efficiency at design speed is improved by 0.06% in spite of extremely limited potential for efficiency enhancement of the original design. The present work is of scientific significance as well as industrial application value in the three-dimensional design optimization of advanced multistage axial-flow compressor at the affordable computational cost.

Performance Evaluation of Transonic Axial Flow Compressor under Distorted Conditions by Groove Casing Technique with Tip Injection and Surface Roughness Effects

The aerodynamic performance and pressure ratio of modern aircraft engine compressor are degrading due to unsteady flow distorted aerodynamic loads and complex phenomena in the blade tip. This paper attempts to improve the performance of modern axial flow transonic compressor stage under distorted flow conditions by incorporating the combination of groove casing, tip injection technique with surface roughness effects. This is performed for the Mach number ranging between 0.8 and 1.3. The obtained numerical results are compared with experimental studies and found to be in good agreement. Through the numerical analysis, the compressor stage flow interaction with the shock waves as well as vortex formation and boundary layer separation is studied in detail. While evaluating the performance of the axial flow compressor, more emphasis is given to the flow properties like pressure, density, temperature, velocity, etc. The performance improvement is observed at a particular Mach number for a specified aerodynamic flow property.

Strength and vibration analysis of transonic compressor under multiple field loads

The transonic axial flow compressor Rotor 37 is subjected to the combined action of thermal load, centrifugal load, and aerodynamic load in real work. The difference in load will cause the blade deformation and the size and distribution of stress to change. And it will cause the vibration frequency of the blade to change compared with the static state at room temperature. This paper studies the compressor blades based on the fluid-thermal-structure coupling method, and comprehensively considers the combined effects of thermal load, centrifugal load and aerodynamic load on the compressor blades. The results of the study show: that when the thermal load acts on the blade alone, it will have less impact on the blade deformation and greater impact on the stress, but when coupled with other loads, it will have a greater impact on the blade deformation and stress. In the case of fluid-thermal-structure coupling, the deformation of the compressor blade will first increase and then decrease with the increase of mass flow, and the change of stress will not decrease with the increase of mass flow. Rotation speed will affect the change trend of blade stress with mass flow. Compared with the static state at room temperature, the natural frequency of the blade under fluid-thermal-structure coupling has undergone a huge change. The first-order vibration frequency has increased by 25.77 %. It can be seen from the Campbell diagram that the compressor blade has a resonance at about 63 % and 98 % Point, in this case, the blade is prone to resonance, and more attention should be paid.

Critical endwall blockage attenuation-based automatic optimization of casing treatment design for transonic axial flow compressor

Tip leakage flow (TLF) and endwall blockage formation are closely associated with the inception of rotating stall in tip-critical transonic axial flow compressor (AFC) rotors. The present study proposes a critical endwall blockage attenuation-based optimization method of casing treatment design for transonic AFC. First, a well-defined endwall blockage quantification method is developed to reveal the development of the endwall blockage at near-stall, and two blockages of the maximum endwall blockage and the near trailing edge endwall blockage are identified as the critical values for stall inception. Next, both adiabatic efficiency at design point and critical endwall blockages at near-stall are used as optimization objectives in an adaptive surrogate-based optimization procedure to deal with the design of nonuniform multiple grooves casing treatment (NMGCT) for the NASA Stage 35, thus avoiding the extremely time-consuming simulations of compressor performance curves. The analysis of the representative optimal designs demonstrates that the NMGCT can act on the tip flow field in a variety of ways, e.g. lowering the maximum endwall blockage can suppress the intensity of the TLF-shock wave interaction, or lowering the near trailing edge endwall blockage is able to control both the trajectory of the TLF closer to the suction surface and the thickness of the suction surface boundary layer. These achievements can help improving the stall margin of the tip-critical transonic AFC. However, in the case of AFC similar to Stage 35, reducing the maximum endwall blockage can achieve greater stall margin improvement and smaller peak efficiency penalty.

Computational Prediction of the Performance Map of a Transonic Axial Flow Compressor

Aviation fuel efficiency is an important target for aviation industry. Aircraft engine compression ratio is a key factor to improve fuel consumption. Compression ratio can be increased using transonic compressor. In this study, performance prediction of a transonic axial compressor at design and off-design operating conditions is investigated numerically using ANSYS-CFX software. The compressor is NASA Rotor 37. Firstly, the performance at design point is predicted, where mesh independence study is performed to determine suitable mesh size. Three-dimensional flow details for meridional plane, blade-to-blade plane and airfoil surface are explored. The design point study successfully captured flow features such as shock waves and flow separation regions. When compared with experimental data, the predicted compressor pressure ratio deviation error is less than 5%. 3D flow details show that shock wave strength increases from hub to tip. The shock wave moves backward as we move from hub to tip indicating that the flow separation covers lesser portion of the blade. Secondly, off-design performance is predicted for various rotational speeds. A simple procedure is utilized to predict surge and choke limits. The predicted compressor map is compared with experimental data and it shows overall root mean square error less than 5%. The success of the method developed in this research make it a viable method to be used in the design phase of transonic compressors to evaluate the effect of design modifications for both design and off-design operating conditions.

Effect of different radial skewed angles of reversed blade-angle slot casing treatment on transonic axial flow compressor stability

Single-channel unsteady numerical simulation was carried out on NASA Rotor 35 to study the influence of radial skewed angle of reversed blade-angle slot casing treatment (RBSCT) on the aerodynamic performance and stall margin of a transonic compressor. Moreover, the influence mechanism is explained by detailed flow field analysis. The radial skewed angles were set to +0°, +30°, +60°, and +75° in the research. The calculated result shows that stall margin improvement (SMI) generated by the slots is increased when the radial skewed angle is gradually increased. The SMI of 17.47% for the slots with +75° radial skewed angle is the biggest among the four RBSCTs. As for the design efficiency improvement (DEI), it is increased first and then decreased with the radial skewed angle increasing. RBSCT with +60°radial skewed angle achieves the greatest DEI of 1.11%. The flow field analysis shows that the radial momentum of the injected and sucked flows is improved with the increasing of radial skewed angle. The improvement can reduce the relative airflow angle of mainstream near the stall condition. Furthermore, the excessive radial skewed angle will increase the flow losses in the slots. Consequently, the design efficiency is reduced at +75°.

Geometry scaling technique and application to aerodynamic redesign of multi-stage transonic axial-flow compressors

Geometry scaling technique, as one of the essential steps in aerodynamic redesign of multi-stage compressor, directly determines whether the new scaled geometry meets the required performance or not. It is also one of the effective measures in unsteady flow simulation of multi-stage compressor to reduce greatly the computational cost by scaling the computational domain from full-annulus passages to partial-annulus passages. The common critical issue arisen from both practices is the geometry dissimilarity in which some certain difference exists between the original and scaled geometries due to a diversity of redesign requirements or model simplifications. However, the extent of how the geometry dissimilarity influences the aerodynamic performance of multi-stage transonic compressor is unclear up to now. Particularly, how to keep the aerodynamic performance similarity in the case of geometry dissimilarity has not been explored yet. Here we propose a geometry scaling technique in which 3D parametric geometry modeling and 1D mean-line analysis are combined to define a series of key redesign criteria in order to keep as minor as possible variations in the aerodynamic performance of multi-stage compressor. The proposed technique is applied to redesign a 3.5 stage transonic axial-flow compressor with two geometry scaling strategies, i.e., adjustable blade/vane numbers and inlet guide vane (IGV)-rotor-stator axial spacing. The key aerodynamic parameters determined by the key redesign criteria are then quantified to verify the feasibility and accuracy of the proposed technique by computational fluid dynamics (CFD) simulations. The results show that by keeping the similarity of key redesign criteria, that is, blade solidity, stagger angle, inlet blade angle, outlet blade angle and thickness to chord distribution, the similarity of key aerodynamic parameters (incidence angle, deviation angle, loss coefficient and aerodynamic blockage) and thus of overall aerodynamic performance (mass flow rate, pressure ratio and efficiency) can be generally guaranteed. The variations of aerodynamic performance in the case of geometry dissimilarity with different blade/vane numbers and IGV-rotor-stator axial spacing are generally within 1%. Compared against the similarity of the other key aerodynamic parameters, a certain dissimilarity of aerodynamic blockage is observed, impacting slightly on the mass flow rate and thus stall margin. The present work is of effective assistance to improve the redesign methodology for multi-stage axial-flow compressors.