Single-stage Transonic Axial Compressor Research Articles

Aerodynamic Optimization of a Single-Stage Transonic Axial Compressor with a Circumferential Feedback Channel

Aerodynamic Optimization of a Single-Stage Transonic Axial Compressor with a Circumferential Feedback Channel

AERODYNAMIC PERFORMANCES OF A SINGLE-STAGE TRANSONIC AXIAL COMPRESSOR USING AN INCLINED CASING GROOVE

ABSTRACT This paper presents the effect of an inclined casing groove on the aerodynamic performance of a single-stage transonic axial compressor, NASA Stage 38, using three-dimensional Reynold-averaged-Navier-Stokes equations with the k-ε turbulence model. The research was carried out to examine the effects of four casing groove parameters: angle, width, depth, and location. Validation of a numerical model for a single-stage transonic axial compressor was conducted to evaluate the computational fluid dynamics method. Most of the simulations showed positive results with an increase in stall margin, adiabatic efficiency, and total pressure ratio, in which the maximum stall margin, adiabatic efficiency, and total pressure ratio can be raised by 87.09%, 0.13%, and 1.57%, respectively, as compared to the smooth casing case.

Casing treatment using oblique slots for a single-stage transonic axial compressor

Abstract This work presents a casing treatment design using oblique slots to improve the stability of a single-stage transonic axial compressor, NASA Stage 37. 36 individual slots are distributed uniformly on the rotors’ shroud. Numerical simulations were performed to find the effects of the casing treatment on the compressor’s stability as well as the adiabatic efficiency and pressure ratio. The oblique slots were found to effectively alleviate the negative influence of the tip leakage flow on the compressor’s performance, thus increasing the stall margin without significant sacrifices in efficiency and pressure rise. Three geometric parameters of the casing treatment were examined. The results showed that, with the oblique slots, the highest stall margin reached 13.59% - a 37% increase from the smooth casing case, with about 0.17% and 0.09% reductions in the peak efficiency and maximum pressure ratio, respectively.

Aerodynamic and structural performances of a single-stage transonic axial compressor with blade fillet radius

PurposeThis study’s investigation aims to clarify the effect of an additional geometry, i.e. a fillet radius, to the blades of a single-stage transonic axial compressor, NASA Stage 37, on its aerodynamic and structural performances.Design/methodology/approachApplying the commercial simulation software and the one-way fluid–structure interaction (FSI) approach, this study first evaluated the simulation results with the experimental data for the aerodynamic performances. Second, this paper compared the structural performances between the models with and without fillets.FindingsThis research analyses the aerodynamic results (i.e. total pressure ratio, adiabatic efficiency, stall margin) and the structural outcomes (i.e. equivalent von Mises stress, total deformation) of the single-stage transonic axial compressor NASA Stage 37.Originality/valueThis paper mentions the influence of blade fillets (i.e. both rotor hub fillet and stator shroud fillet) on the compressor performances (i.e. the aerodynamic and structural performances).

Design Optimization of a Dual-Bleeding Recirculation Channel to Enhance Operating Stability of a Transonic Axial Compressor

The present work performed a comprehensive investigation to find the effects of a dual-bleeding port recirculation channel on the aerodynamic performance of a single-stage transonic axial compressor, NASA Stage 37, and optimized the channel’s configuration to enhance the operating stability of the compressor. The compressor’s performance was examined using three parameters: The stall margin, adiabatic efficiency, and pressure ratio. Steady-state three-dimensional Reynolds-averaged Navier–Stokes analyses were performed to find the flow field and aerodynamic performance. The results showed that the addition of a bleeding channel increased the recirculation channel’s stabilizing effect compared to the single-bleeding channel. Three design variables were selected for optimization through a parametric study, which was carried out to examine the influences of six geometric parameters on the channel’s effectiveness. Surrogate-based design optimization was performed using the particle swarm optimization algorithm coupled with a surrogate model based on the radial basis neural network. The optimal design was found to increase the stall margin by 51.36% compared to the case without the recirculation channel with only 0.55% and 0.28% reductions in the peak adiabatic efficiency and maximum pressure ratio, respectively.

A Novel Surface Parametric Method and Its Application to Aerodynamic Design Optimization of Axial Compressors

A novel parametric control method for the compressor blade, the full-blade surface parametric method, is proposed in this paper. Compared with the traditional parametric method, the method has good surface smoothness and construction convenience while maintaining low-dimensional characteristics, and compared with the semi-blade surface parametric method, the proposed method has a larger degree of geometric deformation freedom and can account for changes in both the suction surface and pressure surface. Compared with the semi-blade surface parametric method, the method only has four more control parameters for each blade, so it does not significantly increase the optimization time. The effectiveness of this novel parametric control method has been verified in the aerodynamic optimization field of compressors by an optimization case of Stage35 (a single-stage transonic axial compressor) under multi-operating conditions. The optimization case has brought the following results: the adiabatic efficiency of the optimized blade at design speed is 1.4% higher than that of the original one and the surge margin 2.9% higher, while at off-design speed, the adiabatic efficiency is improved by 0.6% and the surge margin by 1.3%.

Stability Enhancement of a Single-Stage Transonic Axial Compressor Using Inclined Oblique Slots

A casing treatment using inclined oblique slots (INOS) is proposed to improve the stability of the single-stage transonic axial compressor, NASA Stage 37, during operation. The slots are installed on the casing of the rotor blades. The aerodynamic performance was estimated using three-dimensional steady Reynolds-Averaged Navier-Stokes analysis. The results showed that the slots effectively increased the stall margin of the compressor with slight reductions in the pressure ratio and adiabatic efficiency. Three geometric parameters were tested in a parametric study. A single-objective optimization to maximize the stall margin was carried out using a Genetic Algorithm coupled with a surrogate model created by a radial basis neural network. The optimized design increased the stall margin by 37.1% compared to that of the smooth casing with little impacts on the efficiency and pressure ratio.

Effects of Air Injection on Aerodynamic Performance of a Single-Stage Transonic Axial Compressor

This paper presents a parametric study of stator air injection on aerodynamic performances of a single-stage transonic axial compressor, NASA Stage 37, using three-dimensional Reynolds-averaged Navier- Stokes equations with the k-ε turbulence model and scalable wall functions. The curvature, location and width of the stator injector and the mass flow rate of air injection were tested to find their effects on the aerodynamic performances, such as total pressure ratio, peak adiabatic efficiency, stall margin and stable range extension. The numerical results for adiabatic efficiency and total pressure ratio were validated with experimental data for smooth casing. The results of parametric study show that the aerodynamic performances of the single-stage transonic axial compressor improve greatly the peak adiabatic efficiency compared to the compressor with smooth casing.

Influences of slotted blade on performance and flow structure of a transonic axial compressor

Main reason of compressor instability is boundary layer separation on the surface of blades. As one of flow control methods of the compressor, slotted blade has attracted many researchers’ attention because of its simple geometric structure and remarkable flow control effect. In order to evaluate its availability in the compressor, a type of convergent slot is designed to implement in a single-stage transonic axial compressor. Three configurations, i.e. rotor slot, stator slot and rotor-stator combined slot, are introduced to study the aerodynamic performance of compressor by numerical simulations. Furthermore, flow structures have been analyzed to explain the corresponding mechanism. The results show that overall stability margin of the compressor has been improved by flow control with slotted blade. Behavior of the rotor slot is better than that of the stator slot, but due to mass flow leakage in the slot, peak efficiency and chocking mass flow rate of the compressor are decreased by 1.18% and 3.8% respectively. The low momentum flow on pressure surface is sucked into the jet slot of stator blade, which improves the overall stability margin of 0.63%. The combined scheme with slotted rotor and slotted stator has obtained the best aerodynamic behavior with the increase of the overall stability margin of 2.83%. During the future research, main goal will be improvement of the compressor performance and extension of the mass flow rate range.

Recirculation-groove coupled casing treatment for a transonic axial compressor

This paper investigates a novel casing treatment for an axial compressor that combines a circumferential casing groove with flow recirculation channels. This casing treatment is located in the rotor's tip region and comprises a single shallow circumferential groove with 36 recirculation channels that are distributed equally around the annulus. The effects of the casing treatment on the stability and aerodynamic performance were evaluated numerically on a NASA Stage 37 single-stage transonic axial compressor. The results show that the addition of the recirculation channels improved the gain in stall margin while mitigating the penalty in the adiabatic efficiency and pressure rise compared to a casing groove without recirculation channels. Six geometric parameters were examined in a parametric study. The maximum increase in stall margin was 42.5% with small reductions in the efficiency and pressure ratio. To maximize the stall margin and minimize efficiency loss, two design variables were selected to perform a multi-objective optimization using a multi-objective genetic algorithm coupled with surrogate models using a radial basis neural network. The optimization successfully improved both the objective functions for all representative designs compared to the reference design.

Effects of stator splitter blades on aerodynamic performance of a single-stage transonic axial compressor

Splitter blades located between stator blades in a single-stage axial compressor were proposed and investigated in this work to find their effects on aerodynamic performance and operating stability. Aerodynamic performance of the compressor was evaluated using three-dimensional Reynolds-averaged Navier-Stokes equations using the k-e turbulence model with a scalable wall function. The numerical results for the typical performance parameters without stator splitter blades were validated in comparison with experimental data. The numerical results of a parametric study using four geometric parameters (chord length, coverage angle, height and position) of the stator splitter blades showed that the operational stability of the single-stage axial compressor enhances remarkably using the stator splitter blades. The splitters were effective in suppressing flow separation in the stator domain of the compressor at near-stall condition which affects considerably the aerodynamic performance of the compressor.

Aeromechanic performance of a single-stage transonic axial compressor with recirculation-bleeding channels

ABSTRACT This paper presents a new investigation of the recirculation channel combined with bleeding channels, where the recirculation airflow recirculates air from downstream high-pressure region to an upstream through a recirculation channel located on shroud surface in rotor domain and the bleeding airflow through 36 bleeding channels sited on the recirculation channels ejects airflow to the by-pass channel in a single-stage transonic axial compressor used in turbofan engine. The object of this investigation is to reduce the vortex flow in the tip leakage region by finding the effects of recirculation and bleeding channels design on aeromechanic performance of a single-stage axial compressor, NASA Stage 37. A parametric study of the recirculation and bleeding channels were performed using various geometric parameters related to the locations and shapes of the recirculation and bleeding channels to find the effects of recirculation and bleeding channels design on aeromechanic performance, such as pressure ratio, efficiency, stall margin, blade tip deformation and Von-Mises stress. The numerical results showed an increase in aerodynamic performance and decrease in blade tip deformation with recirculation-bleeding channels, compared to the smooth casing.

Effects of Rotor-Bleeding Airflow on Aerodynamic and Structural Performances of a Single-Stage Transonic Axial Compressor

This paper presents the effects of airflow bleed from the rotor shroud surface between leading and trailing edges on aerodynamic and structural performances of a single-stage transonic axial compressor, NASA stage 37, using three-dimensional Reynolds-averaged Navier–Stokes equations with the k–e turbulence model. A small airflow mass flow rate is bleed throughout a rotor-bleeding ejector which designed by seven parameters: bleeding angle (°) and ejection angles (°), ejection depth (D), bleeding thickness (H), bleeding position from rotor leading edge (L) in flow direction, bleeding ejection curvature (R), and bleeding width contacted on rotor shroud surface (W). The numerical results for aerodynamic performance: total pressure ratio, adiabatic efficiency, and stall margin of a transonic axial compressor were validated with a smooth casing experimental data. A parametric study of seven design parameters of rotor-bleeding ejector above combined with a small ejection mass flow rate in a single-stage transonic axial compressor for aerodynamic and structural performances was performed. The numerical results show that all aerodynamic performance increases with bleeding airflow from rotor shroud surface, the total deformation on rotor tip leading edge in spanwise direction reduces with a very small increasing in Von-Mises stress in a reference-bleeding airflow as compared to the results of smooth casing.

Improvements of performance and stability of a single-stage transonic axial compressor using a combined flow control approach

The performance and the stability of axial compressors are seriously concerned as it's widely used in aero-engines and power machines. A new combined flow control approach was developed previously using blade slot and vortex generator, which was used to reduce flow loss and improve flow stability. The approach was experimentally proved to be very effective in linear compressor cascades. To further evaluate its availability in actual compressors, the new approach is implemented in a single-stage transonic compressor, which introduces a part-span slot into the rotor while a full-span slot and a vortex generator into the stator. CFD simulations are performed to evaluate the gains of compressor performance, and flow structures are analyzed to explain the corresponding mechanisms. Results show great improvements of performance and stability of the compressor. The averaged total pressure ratio and isentropic efficiency improve by 1.82% and 0.88% respectively due to the great reduction of separations in both the rotor and the stator. Because the compressor instability is initiated by the stator corner stall, the delay of the corner separations leads to a great increase of the stable mass-flow range by 26.1%. The compressor flow structures are also well improved because the detrimental backflows and large-scale vortices are mostly removed by the combined approach.

Aerodynamic characterization of a transonic axial flow compressor stage – with asymmetric tip clearance effects

This manuscript discusses the performance of a single-stage transonic axial flow compressor with non-uniform/asymmetric rotor tip clearance. Detailed steady and unsteady experimental measurements are carried out for a compressor with non-uniform clearance over the rotor. Measurements are focused at the peak clearance region of the compressor annulus where stall inception is likely to occur. A sector-based steady-state CFD simulation is performed on the compressor stage with uniform averaged rotor running clearance obtained from the blade growth estimates at design speed using structural analysis. The role of the tip leakage vortex on the stall dynamics of the compressor is elucidated. The stall events/disturbances occurring at every two rotor revolutions are shown using hotwire measurements. Varied flow features of the transonic compressor stage are discussed for different compressor speeds investigated. The level of clearance eccentricity/asymmetry existing in this compressor due to its complex spread does not degrade the overall compressor performance behavior.

Numerical investigations on stator hub initiated stall in a single-stage transonic axial compressor

A new kind of stall inception initiated from the stator hub region was recently identified by experiment in a transonic compressor. To further explore the mechanisms of the new stall inception, the detailed stall evolution is studied in this paper using full-annulus unsteady simulations. The simulations correctly predict several key characteristics observed in the previous experiments: (1) the stall precursor is initiated in the stator hub region; (2) the initial disturbance is axisymmetric; (3) asymmetric rotating disturbance is developed afterwards. The numerical results also illustrate that the stall evolution has two distinct phases: the part span stall and the global stall, which are associated with the axisymmetric and the asymmetric disturbances respectively. The axisymmetric disturbance is caused by the ring-shaped flow separation in the stator hub region, while the asymmetric disturbance is initiated by the breakdown of the symmetry of the ring-shaped separation. For both disturbances, the axial velocity waveforms are in anti-phase at the stator tip and hub region, so they have the 1st-order mode in the span-wise direction. Further discussions on the radial distribution of load indicate that the localized critical load is the key factor leading to the earlier occurrence of flow breakdown in the stator hub region.

정익 슈라우드 공기분사가 단단 천음속 축류압축기의 공력성능에 미치는 영향

본 연구에서는 단단 천음속 축류압축기 정익부 슈라우드에 공기분사기를 설치하는 방안을 제시하고 이 것이 공력성능에 미치는 영향을 분석하기 위해 매개변수 연구를 수행하였다. 분사기의 곡률, 폭, 정익 앞전으로부터의 거리, 원주방향 각도 및 공기분사 비율을 변수로 선정하였고, 각 변수의 변화에 따른 공력성능 영향을 분석하기 위해 삼차원 레이놀즈평균 나비어-스톡스 방정식을 사용한 공력해석을 수행하였다. 매개변수 연구 결과, 정익 분사가 적용된 단단 축류압축기의 공력성능이 향상되었고, 공력성능이 공기분사 비율에 민감한 영향을 받는 것이 확인되었다. 원주방향 각도의 비율이 10%일 때 전압력비와 단열 효율이 가장 크게 향상되었다.

Stall inception and warning in a single-stage transonic axial compressor with axial skewed slot casing treatment

Characteristic changes in the stall inception in a single-stage transonic axial compressor with an axial skewed slot casing treatment were investigated experimentally. A rotating stall occurred intermittently in a compressor with an axial skewed slot, whereas spike-type rotating stalls occurred in the case of smooth casing. The axial skewed slot suppressed stall cell growth and increased the operating range. A mild surge, the frequency of which is the Helmholtz frequency of the compressor system, occurred with the rotating stall. The irregularity in the pressure signals at the slot bottom increased decreasing flow rate. An autocorrelation-based stall warning method was applied to the measured pressure signals. Results estimate and warn against the stall margin in a compressor with an axial skewed slot.

Multi-disciplinary design optimization and performance evaluation of a single stage transonic axial compressor

The multidisciplinary design optimization method, which integrates aerodynamic performance and structural stability, was utilized in the development of a single-stage transonic axial compressor. An approximation model was created using artificial neural network for global optimization within given ranges of variables and several design constraints. The genetic algorithm was used for the exploration of the Pareto front to find the maximum objective function value. The final design was chosen after a second stage gradient-based optimization process to improve the accuracy of the optimization. To validate the design procedure, numerical simulations and compressor tests were carried out to evaluate the aerodynamic performance and safety factor of the optimized compressor. Comparison between numerical optimal results and experimental data are well matched. The optimum shape of the compressor blade is obtained and compared to the baseline design. The proposed optimization framework improves the aerodynamic efficiency and the safety factor.

Research on Multi-Solver Simulation Interface of Aero-Engine Compressor

Establish a CFD simulation interface for compressor aerodynamic analysis based on general CFD software. The simulation interface includes data format, process and interface standard. Data format based on CGNS is built for cross-platform restoring and retrieving mass data. The designing of process references the process of general CFD analysis and the compressor aerodynamic analysis. The interface standard is used to import and the parameters and export them into corresponding profiles according to the demand of different computation modules. Validation results for the simulation interface using a single-stage transonic axial compressor fan ATS-2 shows that the setting process is more convenient and achieves a certain degree of automation.