Single-stage Axial Compressor Research Articles

Scaling Laws at Stall in an Axial Compressor with an Upstream Perturbation

An experimental campaign has been conducted to assess the effect of total pressure inlet distortion on a single stage axial compressor. Different grid sectors have been installed in front of the inlet of the compressor to generate the distortion. Pressure measurements and particle image velocimetry have been used to characterize the impact of the distortion on the flowfield and the performance of the compressor. The broad range of porosity and angular extension of the grid sectors has allowed identifying four different flow regimes characterised by a different response of the compressor. In particular, the low porosity introduces flow regimes which have been never encountered in typical distortion tests restricted to high-porosity grids. In two of these regimes, it has been possible to identify a new scaling of the stall margin versus a new distortion index based on more solid physical ground than already existing indices such as the DC60. The experiments have also shown that for each of these four flow regimes behind the grid different rotating stall features (spike stall or modal stall) appear. A scaling of the stall dynamic has been finally provided thanks to analysis of unsteady casing static pressure signals acquired during the stall transient.

Flow stability analysis on low-speed compressor with swept rotor blades under radial inlet distortions

Flow stability enhancement for axial flow compressors under radially distorted inlet conditions through rotor sweep design is an important issue for designers. To investigate the couped effect of rotor swept and radial inlet distortion on flow stability, steady numerical simulations, theoretical model predictions, and experimental validations were conducted for a low-speed, single-stage axial compressor (TA36) under various intensities of tip total pressure distorted inlet conditions with different swept rotor configurations. The steady pressure rise significantly decreases as the intensity of the distortion increases, whereas the rotor sweep has little effect on the pressure rise characteristic. The maximum error via model prediction is 1.25 %, while for steady numerical calculation, it is 3.89 %. Higher intensities of distortion lead to worse flow stability. Under radially distorted inlet conditions, forward sweep enhances flow stability whereas backward sweep further worsens it. What's more, quantitative analysis shows that the stability enhancement of forward sweep cannot fully counteract the adverse impact of radial distortion on flow stability. Elevated rotor tip loading was identified as the primary cause of instability under tip total pressure distorted inlet conditions. Forward sweep can slightly reduce this loading, whereas backward sweep has the opposite effect. Furthermore, the impact of rotor sweep on rotor tip loading intensifies with increasing distortion intensity. Unsteady flow field analysis was conducted. Which shows that radial distortion and blade sweep both affect flow stability by altering the intensity of tip leakage flow.

Integrated Design of Axial Slot Casing Treatment and Blades in a Low-Speed Single Stage Axial Compressor

Abstract Casing treatment has been used to enhance the stability of axial-flow compressor by extending the stall margin; it also induces strong interaction with compressor blades, as demonstrated in the open literature. Revealing the mechanism of this interaction is a key to fully exploit the potential of the stability-enhancing capability of casing treatment without efficiency penalty. For this purpose, an integrated optimization design of rotor and axial slot casing treatment is implemented for a low-speed axial-flow compressor, and the stall margin is increased by 10.89% without penalty on the peak efficiency. Through detailed analysis, two important driving factors for the improvement of stall margin are illustrated: the suction effect that decreases the axial momentum of tip leakage flow, and the injection effect that increases the axial momentum of main flow. Meanwhile, it is found that the combination of forward-deflection axial slot casing treatment and forward-sweep rotor can maximize the enhancement of the momentum exchange, thereby adequately reducing the blockage region induced by tip leakage flow. More specially, the stator is then redesigned based on the matching mechanism between the stator and rotor with axial slot casing treatment in a stage. This redesign can considerably reduce entropy generation and flow separation in the passage and achieve an improvement of 0.87% in efficiency. The results hence demonstrate the superiority of integrated design of axial slot casing treatment and blades in improving both stall margin and efficiency.

Influences of Root Fillet and Dihedral Angle of Bowed Stator Blade on Aerodynamic Performance of Low-Speed Single-Stage Axial Compressor

Influences of Root Fillet and Dihedral Angle of Bowed Stator Blade on Aerodynamic Performance of Low-Speed Single-Stage Axial Compressor

Effects of rotating inlet distortion on the stall mechanism of an axial flow compressor

To study the influence of rotating inlet distortion on a single-stage axial compressor, both experiment and full-annulus unsteady simulation methods are carried out to reveal the stall mechanism under different inflows. First, the experimental results show that compared with static distortion, rotating distortion reduces the compressor stall margin more seriously. The wavelet analysis of dynamic pressure data shows that although the addition of inlet distortions has changed the stalling process, the compressor still stalls through spike-type. The full-annulus unsteady simulations indicate that the low-velocity zone first appears at the hub under uniform inflow condition. When the compressor is subjected to rotating distortion, the low-velocity zone first appears at the tip region. Time-averaged solutions of unsteady simulations are further studied to reveal the underlying mechanisms. Compared with the flow field under uniform inflow, the addition of rotating distortion can produce co-swirl, but it also causes the change of axial velocity distribution along the radial direction. Under the combined effect of both, the blade loading of rotor tip increases significantly, altering the position where the low-velocity zone first appears.

Investigation of the Stability Criterion for Axial Compressor under Circumferential Total Pressure Distortion

The distorted inlet flow that is a widespread flow phenomenon in compressors not only reduces the compressor performance but more seriously deteriorates the compressor stability, thus threatening flight safety. However, quantitatively determining the stability boundary is challenging due to the complexity of the distorted inlet flow. The effective static-pressure-rise coefficient in the stability criterion from Koch, a stability analysis method under clean inlet flow, was applied to explore the compressor stability under circumferential total pressure distortion. Using a series of single-stage axial compressor experiments for aspect ratio and solidity, this study investigated the influence of the distortion index on compressor stability. Subsequently, based on the analysis method under clean inlet flow, a new stability criterion under circumferential total pressure distortion was established to determine quantitatively the stability boundary. In addition, the stability criterion for determining the stability boundary was combined with the body force model for calculating the compressor performance to simulate a five-stage axial compressor under 120° total pressure distortion, and the results were in agreement with the experimental data to verify the stability criterion. Numerical simulation was used to show the internal flow field in the compressor, analyze the load at each stage, and identify the instability stage that induces stall, which can provide a reasonable solution for determining the stability boundary of a multistage axial compressor under circumferential total pressure distortion. The combination can serve as a convenient preliminary design tool to evaluate compressor stability using fewer computational resources.Practical ApplicationsMost stability investigations adopt simple methods such as numerical divergence and specified parameters under circumferential total pressure distortion, which makes the stability boundary inaccurate. Since the stability analysis methods under clean inlet flow have been studied more deeply so far compared to that under distorted inlet flow, a new stability analysis method under circumferential total pressure distortion is established to quantitatively determine the stability boundary based on the stability analysis method under clean inlet flow. Therefore, in the preliminary design phase, the stability analysis method is combined with the model for calculating the compressor performance to scientifically evaluate the stability of the multi-stage axial compressor under circumferential total pressure distortion. Also, the combination simply requires fewer computational resources than the millions of grids in commercial software.

Effective Inverse Method for High-loaded Axial Compressor Based on Pre-compression Theory

The S1 stream surface inverse method for a single-stage axial transonic compressor was developed and studied under the guidance of the quasi-three-dimensional viscous design theory, based on computational fluid dynamics (CFD), in order to establish and solve governing equations. The re--adjusted load distribution of the S1 stream surface was imposed to obtain a special profile for the rotor, which is known as S-shaping. This changed the structure and weaken the strength of the shock due to the pre-compression effect. In order to match the inlet flow angle of the downstream stator, inverse modification for multiple S1 stream surfaces of the stator blade was conducted in the present study at the same time. As it is known, the S-shaped profile is commonly applied to supersonic flows. Therefore, NASA stage 37 was selected as the design case to verify whether the present inverse method is effective and reliable. Stage 37 was re-designed by stacking five S1 stream surfaces. The profiles of these surfaces were renewed through the inverse design. The results revealed that, compared to the prototype, the aerodynamic performance of the re-designed one was apparently promoted within the stable working range. The adiabatic efficiency at the design point increased by approximately 0.4%, and the pressure ratio improved by approximately 4.5%. In addition, analysis result for the characteristic line revealed that the performance of the redesigned blades at off-design points significantly improved.

Compressor Surges of Near-resonant Type and Convective Type; the Boundaries and the Differences in the Behaviors

Compressor deep surges are observed to be of either near-resonant type or convective type. The differences of the behaviors and the boundaries between both types were tried to make clear in the situations of a single-stage axial compressor on the basis of numerical-experimental results by use of a surge simulation code. In the near-resonant surges, the surge frequencies relative to the resonance frequency range roughly from 0.9 to 1.1, close to the resonance condition. The surges tend to degenerate just above or in the resonance range. The convective type is said to perform actions of emptying and filling the flow in the delivery plenum, showing lower relative surge frequencies less than 0.9. The study for the criteria governing the boundary between both surge types has suggested an axial motion parameter as the candidate. The convective surges are observed to occur above the boundary value of the axial motion parameter. In the environment, the fluctuation amplitudes tend to be larger than those in the near-resonant surges. The situations force the compressor to work over a wider range from the turbine-action zone to the deeply-reverse flow zone along the stage characteristic. The flow reversals thus reinforced will enhance the so-called emptying and filling actions in the convective surges.

Gasdynamic and strength improvement of the axial flow turbine of a helicopter gas turbine engine

The article presents the results of the development of proposals to increase the safety margin of the blades and the efficiency of a single-stage axial flow turbine compressor of a small-sized turbo-shaft engine being modernized with the use of multidiscipline optimization methods. Analysis of the initial version of the turbine showed that the margin of safety of the working blades is significantly lower than that required by the strength standards. To eliminate the problem encountered the task of blade development was initially solved only according to strength characteristics, without taking into account gas dynamic processes. As a result, the geometry of the working blades that met the strength requirements was obtained. It was used as a starting point for solving the joint task of optimizing the blading, taking into account gas dynamic and deformation processes. As a result of joint optimization, a turbine version meeting the strength standards was obtained, although we had to slightly reduce the efficiency of the turbine by 0,2%.

In-Stall Compressor Performance and the Effects of Reynolds Number

Abstract Previous research into axial compressor stall has mainly focused on stall inception and methods to extend the stable operating range. This paper considers the performance of an axial compressor beyond stall and investigates how the characteristics of stall cells depend on Reynolds number. An experimental study has been conducted using a single-stage axial compressor capable of operating across the Reynolds number range of 10,000–100,000. Detailed unsteady measurements have been used to measure the behavior across a range of in-stall flow coefficients. These measurements have been used to extract the stall hysteresis and to determine the size, speed, number, and spanwise extent of the stall cells. The results show that for the stalled compressor, as Reynolds number increases, the size of the minimum stable stall cell decreases. This means that a larger change in throttle area is needed to reduce the stall cell down to a size where the compressor can recover from stall. At Re = 100,000, the stall hysteresis is six times greater than at Re = 20,000. At the design Reynolds number, the number of stall cells that form transitions from one, to two, and then to four stall cells as the flow coefficient is reduced. At lower Reynolds numbers, the two-stall-cell state becomes unstable; instead, a single stall cell transitions directly into five stall cells. In all cases, as the number of stall cells increases, so do the speed of the stall cells and the total size. Further reductions in the flow coefficient cause an increase in the total size of the stall cells and a decrease in the stall cell speed.

Effect of circumferential non-uniform tip clearance on the dynamic stall process of a single-stage axial compressor with total pressure distortion

In this paper, the single-stage compressor with circumferential non-uniform tip clearance is experimentally investigated under 180° total pressure distortion for the compressor characteristics and the dynamic stall process. In the special structure of the circumferential non-uniform tip clearance, different circumferential distortion areas are adopted to actively induce the stall. The maximum or minimum flow coefficient near the stall point occurs when the location where the rotor departs the distortion area is at the average tip clearance rather than the maximum or minimum tip clearance. Based on the time-frequency analysis regarding the dynamic stall process at different correspondences between the inlet distortion and the tip clearance, it is found that the rotating frequency of the stall cell that is independent of the location of the distortion area is slightly less than 50% rotor rotating frequency and the large-scale stall inception whose frequency is 4–8 times the rotor rotating frequency occurs. Besides the circumferential phase difference from 90° to 180° between the location where the disturbance occurs and the location where the rotor departs, the distortion area exists. According to the dynamic stall process, the stall interpretation model of circumferential total pressure distortion under the circumferential non-uniform tip clearance is established.

Stall Inception in a Compressor with Subsonic, Transonic, and Supersonic Inlet Conditions

A combined experimental and computational approach has been used to investigate the stall inception mechanisms and subsequent poststall behavior of a single-stage axial compressor running at different operating conditions, determined by its rotational speed. This paper emphasizes differences in the origin of stall inception depending on the associated inlet flow regimes: 60% speed (subsonic), 80% speed (transonic), and 100% speed (supersonic). At all speeds, stall is initiated with spike-type disturbances that develop into full-span rotating stall. This paper demonstrates for the first time that for a single compressor geometry, a variety of mechanisms can be responsible for the sources of the blockage leading to stall inception depending on the rotational speed. At 60% speed, the inlet flow is subsonic and the primary driver is a corner separation forming between the casing and blade suction surface. At 80% speed, the spillage of tip leakage jet ahead of the leading-edge plane combines with a small shock-induced corner separation to cause sufficient total blockage and subsequent stall inception. Finally, at 100% speed, the mechanisms are the same as at 80% speed but with a greater blockage contribution from the corner separation due to a stronger shock creating a larger separation.

A Method of Stall and Surge Prediction in Axial Compressors Based on Three-Dimensional Body-Force Model

Abstract The occurrence of stall and surge in axial compressors has a great impact on the performance and reliability of aero-engines. Accurate and efficient prediction of the key features during these events has long been the focus of engine design processes. In this paper, a new body-force model that can capture the three-dimensional and unsteady features of stall and surge in compressors at a fraction of time required for URANS computations is proposed. To predict the rotating stall characteristics, the deviation of local airflow angle from the blade surface is calculated locally during the simulation. According to this local deviation, the computational domain is divided into stalled and forward flow regions, and the body-force field is updated accordingly; to predict the surge characteristics, the local airflow direction is used to divide the computational domain into reverse flow regions and forward flow regions. A single-stage axial compressor and a three-stage axial compressor are used to verify the proposed model. The results show that the method is capable of capturing stall and surge characteristics correctly. Compared to the traditional fully three-dimensional URANS method (fRANS), the simulation time for multistage axial compressors is reduced by 1–2 orders of magnitude.

The Effects of Reynolds Number on the Stall and Pre-Stall Behavior of Compact Axial Compressors

Abstract Compact axial compression systems are of interest to the domestic appliance industry. The associated low Reynolds number leads to high losses compared to large-scale compressors due to a transitional flow field with large regions of separation. This paper investigates how Reynolds number variations affect the three-dimensional and unsteady flow field in a compact compressor both pre-stall and in stall. An experimental study has been conducted using a scaled-up single-stage axial compressor across a Reynolds number range of 104–105. Steady and unsteady casing static pressure measurements, along with rotor upstream and downstream unsteady velocity measurements, have been used to observe the rotor flow field. As the Reynolds number is reduced below a critical value, 60,000 in the case of the compressor studied, the pressure rise coefficient of the compressor rapidly decreases. The exact value of the critical Reynolds number is expected to vary with the compressor geometry. This fall-off in performance corresponds to an increase in the compressor rotor secondary flows. Prior to stall, a broadband hump at around 50% of the blade passing frequency is present in the near-field casing static pressure spectra. At Reynolds numbers below the critical value, multiple equally spaced peaks also appear around the peak of the broadband hump. The spacing of these peaks has been found to be exactly equal to the measured stall cell speed once rotating stall is established. When operating in stall, the stall cell is found to increase in circumferential size and slow down as Reynolds number decreases. The measured spectra and observed flow structures show that disturbances exist prior to stall at frequencies consistent with the frequencies within stall. The size and shape of the stall cells that form are related to the extent of the three-dimensional flow field present prior to stall. Below a critical value, all of these flow features are highly sensitive to Reynolds number.

Experimental investigation on axial compressor stall phenomena using aeroacoustics measurements via empirical mode and proper orthogonal decomposition methods

In this work, we conduct experimental investigations on a single-stage axial compressor to shed lights on the dynamic stall phenomena via aeroacoustic measurements. For this, 8 acoustic pressure sensors are installed equally around the circumference of the compressor intake. The acoustic pressure data are simultaneously logged in real-time. Classical and conventional Fourier-transform based methods reveal that the compressor stall will occur beyond a critical pressure ratio or a flow coefficient as illustrated on a compressor map. However, there is no warning precursor obtained from the conventional signal analysis methods. Further investigations are conducted using a number of advanced signal processing techniques, such as empirical mode decomposition (EMD) and proper orthogonal decomposition (POD) and continuous wavelet (CW) methods. EMD analysis reveals that the compressor stall is corresponding to a low-frequency intrinsic mode function (IMF). It is growing rapidly from negligible amplitude random disturbances to limit cycle oscillations. POD shows that the number of the dominant acoustic modes contributing to more than 98.5% of the total fluctuation energy is changed from 3 to 7, when the stall occurs. This is insightful for low-order modeling of the compressor stall phenomena. Finally, applying CW transform could provide a warning of 0.05 s precursor on the tested axial compressor. In general, the present work opens up an alternative approach to study the dynamic physics of a compressor stall by applying an array of acoustic sensors with proper advanced data-processing methods implemented.

The ACONIT project: an innovative design approach of active flow control for surge prevention in gas turbines

The objective of the ACONIT project is to design, manufacture and test actuators for flow control for an implantation in an aircraft engine. The actuators will fulfil aeronautics requirement in order to increase the Technology Readiness Level (TRL) in this domain. In particular, for the present proposal, one plans to focus on the extension of the stable operating range of axial compressor, allowing thus a reduction of the surge margin through postponing the stall onset. To do so, the first objective of the work is to improve the knowledge of the flow physics of an efficient flow control system by joint numerical and experimental analyses performed in a low speed, single stage axial compressor. The results of this analysis will be used to derive the fluidic specifications for high-TRL actuators and control systems. These specifications will be the base for the design and manufacturing of amplified piezo-electric actuator prototypes whose fluidic performance and operational performance in an environment with vibration and controlled level of temperature will be precisely evaluated before manufacturing final actuators that will be integrated in a full-scale engine test facility. Their performance will be evaluated in terms of Surge Margin Improvement (SMI) as well as in terms of energy balance between the induced consumption and the machine performance improvements. The consortium grouped for carrying out this project is composed of a SME (CTEC), two academic institutions (Bundeswehr University Munich and ENSAM) and a Research Centre (ONERA). It groups skills ranging from internal flow analysis in turbomachinery, to flow control or actuators design, manufacturing and characterisations.

Research on the lean and swept optimization of a single stage axial compressor

Compressor design is always a prevalent research issue as nowadays the compressor demands higher performance. Designing afresh a new compressor of higher aerodynamic performance will cost a tremendous amount of time and expense. Consequently, the wiser method of acquiring a higher performance compressor is to optimize a current compressor, discussed in this paper. First, a detailed numerical simulation of al compressor is done. A parametric rotor model is created and the variations of lean and swept parameters are limited. Then, an optimization calculation is carried out based on the combination of an artificial neural network and a genetic algorithm. The optimization iteration is started at three specified working points – a choke point, a design point and a near stall point – to acquire a comprehensive improved solution. A novel penalty function considering both pressure ratio and efficiency is constructed to find the optimal model. Finally, a detailed review is made of the original and the optimized compressor. The computational results show that the isentropic efficiency is successfully increased and the stall margin is remarkably improved. Ultimately, a detailed vortex structure analysis with increasing outlet pressure is discussed. This helps the performance increment mechanism to be understood further.

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.

Numerical investigation of the effects of system volume and average mass flow on the surge characteristics of an axial compressor

The objective of this work is to evaluate the surge characteristics of an axial compressor operating under surge conditions, by using a hybrid BDF/harmonic balance method. During a surge event, a large amount of backflow occurs in the compressor. Consequently, the airflow oscillates along the axial direction at a low frequency, which can considerably damage the compressor. In this paper, a numerical investigation is performed to examine the effects of the system volume and average mass flow on the surge characteristics of a transonic high speed single stage axial compressor (stage 35) designed and evaluated at the NASA Glenn Center. The results show that the oscillation frequency of the airflow in the compressor generally ranges from 14.32–26.04 Hz during surge conditions, and both the system volume and average mass flow considerably influence the surge characteristics. For a constant average mass flow, as the system volume decreases, the oscillation frequency of the airflow in the compressor increases slightly, while the oscillation amplitude of the exit static pressure decreases significantly, and the maximum static pressure at the outlet remains nearly constant. In the case of a constant system volume, with the decrease in the average mass flow, the oscillation frequency of the airflow in the compressor gradually decreases, and the oscillation amplitude of the exit pressure increases significantly. With a further decrease in the average mass flow, the oscillation frequency stabilizes at a certain point and does not change thenceforth. Furthermore, the surge frequency obtained using the numerical method shows a same tendency with the Helmholtz frequency, but there is a considerable difference between them.