Low-speed Research Compressor Research Articles

The loss mechanisms and vortex dynamics of rotor platform step geometries in a low-speed research compressor stage

The construction and assembly of an axial compressor blade row introduces real geometric features, which leads to the deviation of flow characteristics from the intended design. According to the realistic assembly errors observed in the four-stage low-speed research compressor developed by Shanghai Jiao Tong University, two representative step geometries at the rotor platform were abstracted. Unsteady simulations were carried out in the stage environment to study the influence of hub discontinuities on compressor performance and loss mechanisms. The comparisons of smooth hub, elevated hub (EH), and a wedged hub (WH) demonstrated that all the step geometries resulted in increased loss inside both rotor and stator rows. The loss induced by WH was more sensitive to step heights than EH. Under the influence of forward facing step, rotor incidence was reduced and horseshoe vortex (HV) was exacerbated, leading to the accumulation and separation of low momentum fluids at the pressure side corner. This led to a redistribution of losses across the span. In the lower half span, the relative total pressure loss increased almost linearly with step heights, while in the tip region, it reduced slightly. The backward facing step generated the step corner vortices (SCV) that shed and propagated downstream. The SCV interacted with the HV and cavity leakage vortex in the stator passage. Coupled with the augmented stator inflow angle, the corner separation at the suction side was dramatically intensified and led to a significant loss increase.

Optimization and experimental study of stationary endwall of stator labyrinth cavity in a low-speed research compressor

Controlling the leakage flow of the labyrinth seal cavity is very important for the efficiency and aerodynamic matching design of aero-engine high-pressure compressor. However, due to the complexity of the geometry and internal flow of the labyrinth seal cavity, the design optimization of the labyrinth seal cavity is rarely studied. In this paper, a multi-objective design optimization method was used to optimize the stationary endwall of the shrouded stator labyrinth cavity, and the optimal endwall shape was obtained. Based on the optimal geometry, the 3D printing and aerodynamic testing of experimental parts were carried out, and the aerodynamic performance and flow field details of the optimal geometry were analyzed by unsteady numerical simulation. The experimental results show that for the optimal geometry, it has a little effect on the total pressure ratio, but the efficiency at the design point is increased by 1.4 %. The numerical results show that the optimal cavity shape reduces low-energy fluid, thus reduces the total pressure loss of S3. At the near stall point, the optimal S3 still has a better performance, which means that the optimal shape is not limited to the design point. From the result, an engineering design criterion can be concluded. It is suggested that the labyrinth seal cavity should be designed to match the velocity components of the leakage flow with the mainstream of the stator properly to make the leakage flow go evenly in the blade passage instead of gathering to the suction side.

Overall Performance and Loss Analysis of a Low-Speed Research Compressor with Different Nonaxisymmetric Tip Clearance Layouts

Overall Performance and Loss Analysis of a Low-Speed Research Compressor with Different Nonaxisymmetric Tip Clearance Layouts

An improved stall prediction model for axial compressor stage based on diffuser analogy

Stall represents a limit to the useful operation of the axial compressor. Despite its difficulty, a rational and effective stall prediction model is an urgent requirement for compressor designers. In this paper, a stall prediction model for determining the maximum static pressure rise capability of axial flow compressor stage is presented based on Koch's two dimensional diffuser analogy concept. Compared with the original Koch's model, where the flow angles are only used to consider the increase of inlet dynamic pressure factor caused by high stagger angle, this paper takes into account the effect of flow angles on the diffusion length, and derives the mathematical expression of the modified non-dimensional diffusion length. A semi-empirical correlation is presented to account for the effect of camber angle on the stalling static-pressure-rise coefficient of axial compressor stages. By re-evaluating the Reynolds number effect, a blockage indicator is proposed to determine the effect of boundary layer blockage on stalling capability. Good agreement is demonstrated between the predicted and test stalling pressure rise data for a wide range of axial compressor stages and a 4-stage low speed research compressor. Compared with the original Koch's model, the relative error of the modified model in predicting the stall margin is reduced from 22.5% to 3.5%. This model has a guiding significance for the selection of initial design variables in the axial compressor design process.

Some effects of non-axisymmetric tip clearance layout on axial compressor aerodynamics

A steady and unsteady numerical research is carried out to explore some effects of a specific non-axisymmetric tip clearance layout on the overall performance and stability of an axial compressor stage. For a 4-stage low-speed research compressor (LSRC) in Shanghai Jiao Tong University (SJTU), one-eighth annulus of the inlet guide vane and the first stage rotor was modeled for this study. After the validation for the uniform tip clearance case, a specific non-axisymmetric tip clearance layout is chosen from several random cases generated by the Gaussian Probabilistic Density Function method. Unsteady time-averaged results at the near stall condition show that the chosen non-axisymmetric layout can improve the isentropic efficiency by 1.3% and extend the stall margin by 4%. Detailed analyses on flow fields are carried out to interpret the performance improvement. Due to the circumferential layout of clearance sizes, the inlet mass flow and incidence are redistributed in both the radial and circumferential directions. It leads to blade loading and tip leakage flow varying with the tip clearance size. The quantification of blockage manifests that the blockage arising from the tip leakage flow is significantly alleviated in the non-axisymmetric layout, which leads to improvements in overall performance and stall margin. Transient flow fields at the rotor tip are also analyzed at the near stall condition. For the non-axisymmetric layout, low-momentum regions originating from larger clearance sizes oscillate and develop downstream in one blade passage period.

Hub clearance effect in the design space of the cantilevered stator embedded in a 4-stage low-speed research compressor

This paper focuses on the effect of hub clearance in the design space of the highly loaded cantilevered stator. The embedded 1.5 stages of a low-speed research compressor (LSRC) were conducted with Unsteady Reynolds Average Navier-Stokes (URANS) numerical investigation, and the cantilevered stator adopts positive bowed and fore-sweep three-dimensional design. The research details that with the hub clearance increasing from 1.1% to 4.5% span, the loss coefficient and the total leakage momentum of the cantilevered stator correspond to the change of the blade loading near the hub. When designing the inlet metal angle of the rotor downstream the cantilevered stator, emphasis should be given to considering the inter-stage matching below 15% span. The mixing of leakage flow in 1.1% span clearance and 2.5% span clearance is basically completed in the S3 passage, but the mixing of leakage flow in 3.5% span clearance and 4.5% span clearance is still relatively strong downstream of S3. When calculating the relative entropy variation based on Denton’s mixing model, attention should be paid to the relationship between the leakage flow velocity affected by the hub gap and the mainstream velocity, as well as whether the mixing has been completed in the blade passage.

Flow characteristics of a low-solidity cantilevered stator embedded in a 4-stage low-speed research compressor

This research presents the influence of the low-solidity design on the loss and stability of the cantilevered stator and the evolution of the hub leakage vortex. The highly loaded 4-stage low-speed research compressor (LSRC) with an embedded cantilevered stator (CS) of Stator 3 (S3) was experimentally measured at the blade row outlet and in the S3 passage, respectively, and the URANS simulation method was calibrated accordingly. Then, the prototype CS blade number was reduced, namely low-solidity cantilevered stator (LSCS), to increase the blade loading. The diffusion factor ( DF) of LSCS at the midspan at the design working condition is 5.2% higher than that of CS. The URANS results show that LSCS provides higher efficiency for the 1.5 stage compressor with a consistent stall margin. For S3, the total pressure loss of LSCS is lower than CS at the design point (DP) working condition, but it is the opposite at the near stall (NS) working condition. For LSCS at the NS point, the blockage region becomes large and occupies the lower half of the fore blade passage because the low-solidity stator could not provide the enough flow turning ability, and the core of the hub leakage vortex (HLV) moves forward and the intensity is strengthened. The high loss region of LSCS at the near stall condition is consistent with the size of two hub leakage vortices. The first HLV breakdown is caused by unsteady mixing flow, and then it twists downstream and contacts with the trailing edge of the adjacent blade pressure surface. The study of the hub leakage flow characteristics of the low-solidity cantilevered stator can help designers to control flow better.

Flow characteristics on a 4-stage low-speed research compressor with a cantilevered stator

This paper presents the experimental and numerical investigation of flow characteristics on the highly loaded 4-stage low-speed research compressor (LSRC) with a Cantilevered Stator of Shanghai Jiao Tong University. A new three-dimensional design for the cantilevered stator in the LSRC was derived using an asymmetrical circumferential bow, and the cantilevered stator near the hub adopted the fore-loading design. The unsteady Reynolds averaged Navier-Stokes (URANS) numerical results were compared with experimental data and the differences were introduced. The evolution of the hub leakage flow at the near stall (NS) operating condition compared with the design point (DP) was analyzed, and design suggestions were provided for the two operating conditions. The present study details the effects of the hub leakage flow on the highly loaded LSRC, and provides some quantified data between the aerodynamic performance and hub leakage flow. Designers needing to better understand the hub leakage flow of a cantilevered stator in a multi-stage compressor will benefit from this paper.

Periodic Unsteady Tip Clearance Vortex Development in a Low-Speed Axial Research Compressor at Different Tip Clearances

Since the early work on axial compressors, the penalties due to radial clearances between blades and side walls are known and are an ongoing focus of research work. The periodic unsteadiness of the tip clearance vortex (TCV), due to its interaction with the stator wakes, has only rarely been addressed in research papers so far. The current work presents experimental and numerical results from a four-stage low-speed research compressor (LSRC) modeling a state-of-the-art compressor design. Time-resolved experimental measurements have been carried out at three different rotor tip clearances (gap to tip chord: 1.5%, 2.2%, 3.7%) to cover the third rotor's casing static pressure and exit flow field. These results are compared with either steady simulations using different turbulence models or harmonic Reynolds-averaged Navier–Stokes (RANS) calculations to discuss the periodical unsteady TCV development at different clearance heights. The prediction of the local tip leakage flow is clearly improved by the explicit algebraic Reynolds stress model (EARSM) turbulence model compared to the standard shear stress transport (SST) model. The harmonic RANS calculations (using the SST model) improve the prediction of time-averaged pressure rise and are used to analyze the rotor stator interaction in detail. The interaction of the rotor tip flow field with the passing stator wakes causes a segmentation of the TCV and results in a sinusoidal variation in blockage downstream the rotor row.

Effect of Injection Processing Parameters on Residual Stress for Compressor Experimental Blade Using FEM and RSM

Injection molding-induced residual stresses have strong influence on molding precision for compressor experimental blade used on LSRC (Low Speed Research Compressor). In this paper, finite element model (FEM) and response surface methodology (RSM) are employed to reveal the effect of processing parameters on residual stress for injection molding of compressor experimental blade. First, the influence of processing parameters such as melt temperature, mold temperature on the residual stress of blade was studied based on factorial experiment. Then, the rules of influence of the main processing parameters and their interactions on injection residual stress of blade was analyed based on RSM ． The results show that the packing time and melt temperature are the principle factors for controlling injection residual stress of compressor experimental blade ． In addition, the packing time and melt temperature has strong interaction effect on the residual stress of blade.

Investigation of Performance and Rotor Tip Flow Field in a Low Speed Research Compressor with Circumferential Groove Casing Treatment at Varying Tip Clearance

Experimental investigations in a single-stage low speed axial research compressor are presented. The influence of four different rotor tip clearances on the overall compressor performance and on the rotor tip flow field is investigated in configurations with and without circumferential groove casing treatments. Piezo-resistive pressure transducers are used to capture the unsteady flow field in the rotor tip region. The investigated casing groove is effectively working at the three largest investigated tip clearance sizes. The largest achieved operating range increase by the groove is 6.9%. The groove can delay the upstream movement of the flow interface between leakage flow and main flow and thus increase the stable operating range. Rotating instabilities are shown to exist at large tip clearance sizes in throttled operating conditions. Their amplitudes can be damped by the casing groove. No modal activities could be detected in the current single-stage compressor build.

Modeling Nonuniform Bleed in Axial Compressors

The coupling between the bleed system and the flowfield of a downstream compressor stage is studied using two approaches. In the first approach, three-dimensional, full annulus, unsteady computations simulate the flow in a low-speed research compressor with nonuniform bleed extraction. Comparisons with experimental data show that the flow prediction in the main annulus is accurate to within 0.005 of flow coefficient and 0.5deg of flow angle. The computational fluid dynamics (CFD) is then used to provide a description of flow within the bleed system itself. In the second approach, a two-dimensional mean radius model, similar to that adopted by Hynes and Greitzer in the previous work on compressor stability, is used to simulate the response of the compressor to nonuniform bleed. This model is validated against experimental data for a single-stage compressor, and despite the inherent assumptions (two-dimensional flow and simplified compressor response), provides a satisfactory prediction of the flow for preliminary design purposes with orders of magnitude less computational cost than full 3D CFD. The model is then used to investigate the effect of different levels of bleed nonuniformity and of varying the axial distance between the bleed and the downstream stage. Reducing bleed nonuniformity and moving the stage away from the bleed slot are predicted to reduce the circumferential nonuniformity of the flow entering the stage.

Effects of probe support on the stall characteristics of a low-speed axial compressor

In order to investigate the effects of probe support on the stall characteristics of micro compressors, an experiment was carried out on a large-scale low-speed research compressor according to the principle of geometric similarity. A cylindrical probe support intruding to 50% blade span was mounted at 50% chord upstream from the rotor blade leading edge. The static pressure rise characteristic of the compressor is measured, with and without the probe support respectively. The dynamic compressor behavior from pre-stall to full stall was also measured. The results indicate that the stability margin of the compressor is lowered after installing the probe support. The stall inception is aroused by modal wave disturbances. The disturbances developed into two stall cells smoothly before installing the probe support, while the disturbances first developed into a single stall cell then splitting into two stall cells after installing the probe support. The presence of probe support lowers the initial intensity of the rotating stall of the compressor, while it doesn’t alter the intensity of the rotation stall after the compressor enters into full stall.

Investigation of the internal flow field for a subsonic stage of high pressure compressor through model testing

A large-scale four-stage low-speed research compressor was designed to represent the flow field structure in a subsonic stage, which is the seventh stage of a high pressure compressor. The aim of the design work was to undertake a careful high-to-low speed transformation. The low-speed compressor was designed as an “approximate repeating-stage” configuration, in which the third stage is the model stage and is slightly different from the other stages, while the first two stages and the last stage set up inlet and outlet conditions for it, respectively. The tip–hub ratio, solidity, and aspect ratio of the low-speed research compressor were restricted to be as close as possible to those of high pressure compressor. The blade sections along the span were designed to match the distributions of high-speed surface static pressure coefficient; furthermore, stacking rules of high pressure compressor apply to low-speed blades. The low-speed blades have been tested at the 4-Stage Research Compressor Facility of Nanjing University of Aeronautics and Astronautics. Detailed traverse flow measurements were conducted in the stator blade passages of the third stage and between blade rows, along with static pressure measurements on stator blade surface and particle image velocimetry measurements in the rotor passage. Experiment results show that due to the boundary layer separated below the mid-span of the rotor blade, the total pressure loss increases at the hub region of both the rotor and stator. In addition, the experimental results in the stator blade passage reflect the characteristics of the flow field in the bowed stator passage. The results indicated that the seventh stage of high pressure compressor needs to be optimized by inhibiting the boundary layer separation mainly on the rotor blade surface below the mid-span, which is our further work.

Effects of probe support on the flow field of a low-speed axial compressor

This paper presents an investigation on the effect of probe support on the flow field of an axial compressor. The experiment is carried out in a large-scale low-speed research compressor. A cylindrical probe support intruding to 50% blade span was installed at 50% chord upstream from the rotor leading edge. The region from 5° to 32° off the probe support in the direction of rotation at the rotor outlet was measured with a 5-hole probe and a high-response total pressure probe. The experiment is performed at both near-design and near-stall points. The measuring results of 5-hole probe and high-response total pressure probe indicate that the probe blockage effect is different at different blade spans. The wake of the probe support weakens the leakage vortex intensity at the tip region, leading to greater total pressure rise. At near-design condition, the presence of probe support has a negative effect on the region from 75% to 92% span, while improves the flow field below 75% span. At near stall condition, the probe support has a negative effect on the region from 70% to 90% span, and almost has no influence on the flow field below 70% span.

Design Work of a Compressor Stage Through High-To-Low Speed Compressor Transformation

Low-speed model testing has advantages such as great accuracy and low cost and risk, so it is widely used in the design procedure of the high pressure compressor (HPC) exit stage. The low-speed model testing project is conducted in Nanjing University of Aeronautics and Astronautics (NUAA) to represent aerodynamic load and flow field structure of the seventh stage of a high-performance ten-stage high-pressure compressor. This paper outlines the design work of the low speed four-stage axial compressor, the third stage of which is the testing stage. The first two stages and the last stage provide the compressor with entrance and exit conditions, respectively. The high-to-low speed transformation process involves both geometric and aerodynamic considerations. Accurate similarities demand the same Mach number and Reynolds number, which will not be maintained due to motor power/size and its low-speed feature. Compromises of constraints are obvious. Modeling principles are presented in high-to-low speed transformation. Design work was carried out based on these principles. Four main procedures were conducted successively in the general design, including establishment of low-speed modeling target, global parameter design of modeling stage, throughflow aerodynamic design, and blading design. In global parameter design procedure, rotational speed, shroud diameter, hub-tip ratio, midspan chord, and axial spacing between stages were determined by geometrical modeling principles. During the throughflow design process, radial distributions of aerodynamic parameters such as D-factor, pressure-rise coefficient, loss coefficients, stage reaction, and other parameters were obtained by determined aerodynamic modeling principles. Finally, rotor and stator blade profiles of the low speed research compressor (LSRC) at seven span locations were adjusted to make sure that blade surface pressure coefficients agree well with that of the HPC. Three-dimensional flow calculations were performed on the low-speed four-stage axial compressor, and the resultant flow field structures agree well with that of the HPC. It is worth noting that a large separation zone appears in both suction surfaces of LSRC and HPC. How to diminish it through 3D blading design in the LSRC test rig is our further work.

Tip Flow Unsteadiness and Blade Row Interactions for a Low-Speed Compressor Rotor

A numerical investigation has been conducted to determine the effects of self-excited tip-leakage flow unsteadiness, upstream stator wakes, and downstream blade row interactions on the performance prediction of a low-speed research compressor rotor. Calculations included a single blade row rotor configuration as well as two multiblade row configurations: one in which the rotor was modeled with an upstream stator and a second in which the rotor was modeled with a downstream stator. Steady-state and time accurate calculations were performed using a Reynolds averaged Navier–Stokes solver, and the results were compared with detailed experimental data at operating conditions corresponding to the compressor design point, near stall, and a midpoint between the two. Differences in the performance predictions between the three configurations were then used to determine the effect of the upstream stator wakes and the downstream blade row interactions. Results obtained show that for this compressor with rather large tip clearance time accurate calculations are a vast improvement over steady-state calculations, but the upstream stator wakes and the downstream blade row interactions have only a small effect on the numerical solution. In this paper, changes in the rotor flowfield due to the upstream and downstream stators are investigated and discussed to explain the limited impact of the stator wakes and downstream blade row interactions on this tip dominated flow. In addition, the unsteady nature of the rotor tip vortex at near stall conditions is discussed to explain the improvement of time accurate analysis over steady-state computations.

Effect of Rotor Tip Gap Variation at the Rear Stages of an Axial Flow Compressor

In this paper, detailed analysis of the tip flow of an axial compressor rotor blade has been carried out using the commercial CFD package ANSYS CFX. The rotor blade was designed such that it is reminiscent of the rear stages of a multi-stage axial compressor. The effects of varying tip gaps are studied using CFD simulations for overall pressure rise and flow physics of the tip flow at the design point and near the peak pressure point. Rig tests of a low speed research compressor rotor with 3% tip clearance provided characteristics plots for validation of the CFD results. With increase in clearance from 1% to 4%, the rotor pressure rise at the design point was observed to decrease linearly. Increase in the clearance increases the cross flow across the tip; however, the magnitude of the average jet velocity crossing the tip decreases. The tip leakage vortex was observed to stay close to the suction surface with increase in clearance.

Active flow control at a 1.5-stage low-speed research compressor with varying rotor tip clearance

The stable operating range of axial compressors is limited by the onset of rotating stall and surge. Mass injection upstream of the tip of an axial compressor rotor is a stability enhancement approach which can be effective in suppressing stall in tip-critical rotors, and thus increasing the operating range of compressors. In this article, investigations on active flow control related to the rotor tip gap sensitivity are discussed. The experiments were performed in a 1.5-stage low-speed research compressor. Measurements at part speed (80 per cent) and full speed (100 per cent) with varying injection rates are discussed. These tests were performed for two rotor tip clearances of 1.3 per cent and 4.3 per cent of rotor blade tip chord. Results on the compressor map, the flow field as well as transient measurements to identify the stall inception are discussed. Supplementary, the numerical results are compared to the experiments based on the configuration with the greatest benefit in operating range enhancement.

Numerical investigation of the influence of non-axisymmetric hub contouring on the performance of a shrouded axial compressor stator

This numerical study uses a two-stage low speed research compressor with a shrouded stator row to investigate the interaction of non-axisymmetric endwall contouring at the stator hub with the leakage flow emanating from the shroud cavity. It can be shown that the contour is effective in reducing the hub corner stall caused by the separating endwall boundary layer. The leakage flow itself is not reduced. Subsequently, the height of the shroud seal fins is varied. A comparison of the results with and without non-axisymmetric endwalls shows that the effectiveness of the endwall contour in improving stator performance increases with the amount of leakage flow. Consequently, the sensitivity towards shroud leakage is decreased with contoured endwall. This leads to the conclusion that this design feature does not only serve to improve the flow quality in a given case, but also provides a means to the aerodynamic designer to introduce more robustness into the compressor.