Numerical Investigation of Axial Compressor Stages With Differing Degrees of Reaction

2014 ◽  
Author(s):  
Jens Ortmanns
Keyword(s):  
Numerical Investigation ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Pressure Rise ◽  
Wall Region ◽  
Compressor Stage ◽  
Tip Leakage Flow ◽  
Degree Of Reaction ◽  
Multi Stage ◽  
Flow Phenomena

In order to increase the efficiency of a compressor module, several loss sources such as aerofoil profile loss, secondary loss and clearance flow phenomena must be taken into account and balanced in the most efficient way. This current document presents the results of a numerical investigation based on a conventionally loaded high pressure compressor stage with different inlet and exit swirls. The effects of changing the degree of reaction on the compressor stage flow pattern is analysed in detail. In general, the correlation between the overall stage efficiency at constant pressure ratio and the degree of stage reaction is low. Nevertheless, the results show a direct impact on the rotor tip leakage flow and the secondary flow phenomena in the stator end-wall region when the degree of reaction is modified which is driven by the change in static pressure rise between the rotor and the stator passages. The balance of these two loss sources might have an impact on the efficiency and the stall behaviour of a multi-stage compressor.

Multi Stage Axial Compressor Design and Performance Evaluation

2011 ◽  
Author(s):  
Young Seok Kang ◽  
Tae Choon Park ◽  
Oh Sik Hwang ◽  
Soo Seok Yang
Keyword(s):  
Performance Test ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Pressure Rise ◽  
Multi Stage ◽  
Compressor Design ◽  
Test Result ◽  
Highly Loaded ◽  
Small Aircraft

Recently, needs for Unmanned Air Vehicle (UAV) and small aircraft are increasing and demands for small turbo jet or turbo fan engines are also increasing. Then, size and weight are the two main restrictions in UAV or small aircraft propulsion system applications. One method for resolving such a problem is to increase the pressure rise per stage and to reduce the number of stages. Nowadays, matured compressor aerodynamic design techniques enable us to design highly loaded axial compressors. This paper covers from the design step of a highly loaded transonic axial compressor to the performance test result and its analysis. At the fore part of the paper, aerodynamic process of a multi stage axial compressor is introduced. To satisfy both of the mass flow and pressure rise, the compressor should rotate at a high rotational speed. Therefore the transonic flow field forms in the rotor stages and it is designed with a relatively high pressure rise per stage to satisfy its design target. Basically, one dimensional and quasi three dimensional compressor design were carried with compressor design codes. The compressor stage consists of 3 stages, and the bulk pressure ratio is 2.5. The first stage is burdened with the highest pressure ratio and less pressure rises occur in the following stages. Also it is designed that tip Mach number of the first rotor row does not exceed 1.3. The final design was confirmed by iterating three dimensional CFD calculations to satisfy design target and some design intentions. In the latter part of the paper, its performance test processes are briefly introduced. The performance test result showed that the overall compressor performance targets; pressure ratio and efficiency are well achieved. From the test results, we found some clues for further improvement and optimization of the compressor aerodynamic performance.

The Effects of Tip Leakage Flow on the Performance of Multistage Compressors Used in Small Core Engine Applications

2015 ◽  
Vol 138 (5) ◽  
Author(s):  
Reid A. Berdanier ◽  
Nicole L. Key
Keyword(s):  
Experimental Data ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Pressure Rise ◽  
Tip Clearance ◽  
Tip Leakage Flow ◽  
Stall Margin ◽  
Tip Leakage ◽  
Small Core ◽  
Thermal Growth

Large rotor tip clearances and the associated tip leakage flows are known to have a significant effect on overall compressor performance. However, detailed experimental data reflecting these effects for a multistage compressor are limited in the open literature. As design trends lead to increased overall compressor pressure ratio for thermal efficiency benefits and increased bypass ratios for propulsive benefits, the rear stages of the high-pressure compressor will become physically small. Because rotor tip clearances cannot scale exactly with blade size due to the margin needed for thermal growth considerations, relatively large tip clearances will be a reality for these rear stages. Experimental data have been collected from a three-stage axial compressor to assess performance with three-tip clearance heights representative of current and future small core machines. Trends of overall pressure rise, stall margin, and efficiency are evaluated using clearance derivatives, and the summarized data presented here begin to narrow the margin of tip clearance sensitivities outlined by previous studies in an effort to inform future compressor designs. Furthermore, interstage measurements show stage matching changes and highlight specific differences in the performance of rotor 1 and stator 2 compared to other blade rows in the machine.

Modeling of Axial Compressor With Large Tip Clearances

2020 ◽  
Author(s):  
Simon Evans ◽  
Junsok Yi ◽  
Sean Nolan ◽  
Liselle Joseph ◽  
Michael Ni ◽  
...  
Keyword(s):  
Technology Development ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Pressure Rise ◽  
Core Size ◽  
Detached Eddy Simulation ◽  
Tip Leakage Flow ◽  
Mesh Topology ◽  
Radial Profiles ◽  
Predictive Design

Abstract In the drive for lower fuel consumption through increased bypass ratio and increased overall pressure ratio (OPR), engine designs for the next generation of single-aisle aircraft will require core sizes below 3 lb/s and OPRs above 50. Traditionally, these core sizes are the domain of centrifugal compressors, but materials limit pressure ratio in these machines to well below 50. An all-axial high pressure compressor (HPC) at this core size, however, comes with limitations associated with the small blade spans at the back of the HPC, as clearances, fillets and leading edges do not scale with the core size. The result is a substantial efficiency penalty, driven primarily by the tip leakage flow produced by the larger clearance-to-span ratio, which negates the cycle efficiency benefits of the high OPR. In order to enable small-core, high-OPR, all-axial compressors mitigating technologies need to be developed and implemented to reduce the large clearance-to-span efficiency penalty. However, for this technology development to be successful, it is imperative that predictive design tools accurately model the overall flow physics and trends of the technologies developed. In this paper we describe an effort to determine whether different modeling standards are required for a large clearance-to-span ratio, and if so, identify criteria for an appropriate solver and/or mesh. Multiple models are run and results compared with data collected in the NASA Glenn Research Center’s Low-Speed Axial Compressor. These comparisons show that steady Reynolds-Averaged Navier-Stokes (RANS) solvers can predict the pressure-rise characteristic to an acceptable level of accuracy, if careful attention is paid to mesh topology in the tip region. However, unsteady tools are necessary to accurately capture radial profiles of blockage and total pressure. The Delayed-Detached Eddy Simulation model was also used to run this geometry, but did not resolve any additional features not captured by the unsteady RANS simulation near stall.

Studies on Downstream Stator With Rotor Re-Staggering and Forward Sweeping in a Subsonic Axial Compressor Stage

2011 ◽  
Author(s):  
P. V. Ramakrishna ◽  
M. Govardhan
Keyword(s):  
Flow Field ◽  
Computational Model ◽  
Flow Rate ◽  
Total Pressure ◽  
Axial Compressor ◽  
Pressure Rise ◽  
Axial Distance ◽  
Compressor Stage ◽  
Numerical Work ◽  
Stagger Angle

The present numerical work studies the flow field in subsonic axial compressor stator passages for: (a) preceding rotor sweep (b) preceding rotor re-staggering (three stagger angle changes: 0°, +3° and +5°); and (c) stator sweeping (two 20° forward sweep schemes). The following are the motives for the study: at the off-design conditions, compressor rotors are re-staggered to alleviate the stage mismatching by adjusting the rows to the operating flow incidence. Fundamental to this is the understanding of the effects of rotor re-staggering on the downstream component. Secondly, sweeping the rotor stages alters the axial distance between the successive rotor-stator stages and necessitates that the stator vanes must also be swept. To the best of the author’s knowledge, stator sweeping to suit such scenarios has not been reported. The computational model for the study utilizes well resolved hexahedral grids. A commercial CFD package ANSYS® CFX 11.0 was used with standard k-ω turbulence model for the simulations. CFD results were well validated with experiments. The following observations were made: (1) When the rotor passage is closed by re-staggering, with the same mass flow rate and the same stator passage area, stators were subjected to negative incidences. (2) Effect of stator sweeping on the upstream rotor flow field is insignificant. Comparison of total pressure rise carried by the downstream stators suggests that an appropriate redesign of stator is essential to match with the swept rotors. (3) While sweeping the stator is not recommended, axial sweeping is preferable over true sweeping when it is necessary.

Assessment of a Novel Non-Axial Counter-Rotating Compressor Concept for Aero-Engines

2018 ◽  
Author(s):  
Quentin Dejour ◽  
Huu Duc Vo
Keyword(s):  
Pressure Rise ◽  
Axial Direction ◽  
Boundary Layer Separation ◽  
Low Flow ◽  
Compressor Stage ◽  
Mixed Flow ◽  
Stall Margin ◽  
High Loss ◽  
Multi Stage ◽  
Flow Compressor

This paper presents the first assessment of a new non-axial counter-rotating compressor concept. This concept consists of replacing the stator of a mixed-flow compressor stage or the diffuser of a centrifugal compressor stage with a counter-rotating rotor that will turn the flow back to the axial direction with much lower diffusion factor, while providing the equivalent in work of the upstream mixed-flow rotor or impeller. This concept has two advantages. First, the very high stage pressure rise means that only a single counter-rotating rotor may be required, making mechanical implementation simpler than for multi-stage axial counter-rotating compressors. Second, the replacement of the high flow turning (high loss) stator/diffuser in a non-axial stage with a low flow turning counter-rotating rotor gives the new concept potential for achieving higher efficiency than conventional non-axial compressors. As a first proof of concept, a subsonic counter-rotating mixed-flow compressor and its conventional (i.e. rotor-stator) equivalent have been designed with the intent of being implemented in a test rig. CFD simulations have been carried out for a comparative evaluation of both configurations. Results show that the counter-rotating mixed-flow compressor produces more than double the pressure rise of its conventional version with a slightly higher peak-efficiency while having a smaller axial length. Moreover, the counter-rotating configuration has a better stall margin than its conventional counterpart, for which the boundary layer separation from excessive flow turning in the stator causes early stall.

Performance of Strongly Bowed Stators in a Four-Stage High-Speed Compressor

2004 ◽  
Vol 126 (3) ◽  
pp. 333-338 ◽  
Author(s):  
Axel Fischer ◽  
Walter Riess ◽  
Joerg R. Seume
Keyword(s):  
Mass Flow ◽  
High Speed ◽  
Static Pressure ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Pressure Rise ◽  
Maximum Pressure ◽  
Controlled Diffusion ◽  
Stage 4 ◽  
The University

The FVV sponsored project “Bow Blading” (cf. acknowledgments) at the Turbomachinery Laboratory of the University of Hannover addresses the effect of strongly bowed stator vanes on the flow field in a four-stage high-speed axial compressor with controlled diffusion airfoil (CDA) blading. The compressor is equipped with more strongly bowed vanes than have previously been reported in the literature. The performance map of the present compressor is being investigated experimentally and numerically. The results show that the pressure ratio and the efficiency at the design point and at the choke limit are reduced by the increase in friction losses on the surface of the bowed vanes, whose surface area is greater than that of the reference (CDA) vanes. The mass flow at the choke limit decreases for the same reason. Because of the change in the radial distribution of axial velocity, pressure rise shifts from stage 3 to stage 4 between the choke limit and maximum pressure ratio. Beyond the point of maximum pressure ratio, this effect is not distinguishable from the reduction of separation by the bow of the vanes. Experimental results show that in cases of high aerodynamic loading, i.e., between maximum pressure ratio and the stall limit, separation is reduced in the bowed stator vanes so that the stagnation pressure ratio and efficiency are increased by the change to bowed stators. It is shown that the reduction of separation with bowed vanes leads to a increase of static pressure rise towards lower mass flow so that the present bow bladed compressor achieves higher static pressure ratios at the stall limit.

Effect of Airfoil Vortex Generator on the Performance and Stability of a Transonic Axial Compressor

2021 ◽  
Author(s):  
Subbaramu Shivaramaiah ◽  
Mahesh K. Varpe
Keyword(s):  
Pressure Ratio ◽  
Research Work ◽  
Vortex Generator ◽  
Leading Edge ◽  
Leakage Flow ◽  
Compressor Stage ◽  
Tip Leakage Flow ◽  
Stall Margin ◽  
Tip Leakage ◽  
Transonic Compressor

Abstract In the present research work, effect of airfoil vortex generator on performance and stability of transonic compressor stage is investigated through CFD simulations. In turbomachines vortex generators are used to energize boundary and generated vortex is made to interact with tip leakage flow and secondary flow vortices formed in rotor and stator blade passage. In the present numerical investigation symmetrical airfoil vortex generator is placed on rotor casing surface close to leading edge, anticipating that vortex generated will be able to disturb tip leakage flow and its interaction with rotor passage core flow. Six different vortex generator configuration are investigated by varying distance between vortex generator trailing edge and rotor leading edge. Particular vortex generator configuration shows maximum improvement of stall margin and operating range by 5.5% and 76.75% respectively. Presence of vortex generator alters flow blockage by modifying flow field in rotor tip region and hence contributes to enhancement of stall margin. As a negative effect, interaction of vortex generator vortices and casing causes surface friction and high entropy generation. As a result compressor stage pressure ratio and efficiency decreases.

Flow Characteristics of a Novel Centrifugal Compressor Design

2016 ◽  
Author(s):  
Shashank Mishra ◽  
Shaaban Abdallah ◽  
Mark Turner
Keyword(s):  
Centrifugal Compressor ◽  
Pressure Ratio ◽  
Flow Characteristics ◽  
Axial Compressor ◽  
Low Pressure ◽  
Flow Path ◽  
Optimization Techniques ◽  
Single Stage ◽  
Multi Stage ◽  
Compressor Design

Multistage axial compressor has an advantage of lower stage loading as compared to a single stage. Several stages with low pressure ratio are linked together which allows for multiplication of pressure to generate high pressure ratio in an axial compressor. Since each stage has low pressure ratio they operate at a higher efficiency and the efficiency of multi-stage axial compressor as a whole is very high. Although, single stage centrifugal compressor has higher pressure ratio compared with an axial compressor but multistage centrifugal compressors are not as efficient because the flow has to be turned from radial at outlet to axial at inlet for each stage. The present study explores the advantages of extending the axial compressor efficient flow path that consist of rotor stator stages to the centrifugal compressor stage. In this invention, two rotating rows of blades are mounted on the same impeller disk, separated by a stator blade row attached to the casing. A certain amount of turning can be achieved through a single stage centrifugal compressor before flow starts separating, thus dividing it into multiple stages would be advantageous as it would allow for more flow turning. Also the individual stage now operate with low pressure ratio and high efficiency resulting into an overall increase in pressure ratio and efficiency. The baseline is derived from the NASA low speed centrifugal compressor design which is a 55 degree backward swept impeller. Flow characteristics of the novel multistage design are compared with a single stage centrifugal compressor. The flow path of the baseline and multi-stage compressor are created using 3DBGB tool and DAKOTA is used to optimize the performance of baseline as well novel design. The optimization techniques used are Genetic algorithm followed by Numerical Gradient method. The optimization resulted into improvements in incidence and geometry which significantly improved the performance over baseline compressor design. The multistage compressor is more efficient with a higher pressure ratio compared with the base line design for the same work input and initial conditions.

Design and Test of a Novel Highly-Loaded Compressor

2019 ◽  
Author(s):  
Chengwu Yang ◽  
Ge Han ◽  
Shengfeng Zhao ◽  
Xingen Lu ◽  
Yanfeng Zhang ◽  
...  
Keyword(s):  
Pressure Ratio ◽  
Axial Compressor ◽  
Pressure Rise ◽  
Tip Clearance ◽  
Leakage Flow ◽  
Flow Angle ◽  
Stall Margin ◽  
High Pressure Ratio ◽  
Numerical Predictions ◽  
Highly Loaded

Abstract The blades of rear stages in small size core compressors are reduced to shorter than 20 mm or even less due to overall high pressure ratio. The growing of tip clearance-to-blade height ratio of the rear stages enhance the leakage flow and increase the possibility of a strong clearance sensitivity, thus limiting the compressor efficiency and stability. A new concept of compressor, namely diffuser passage compressor (DP), for small size core compressors was introduced. The design aims at making the compressors robust to tip clearance leakage flow by reducing pressure difference between pressure and suction surfaces. To validate the concept, the second stage of a two-stage highly loaded axial compressor was designed with DP rotor according to a diffuser map. The diffuser passage stage has the same inlet condition and loading as the conventional compressor (CNV) stage, of which the work coefficient is around 0.37. The predicted performance and flow field of the DP were compared with the conventional axial compressor in detail. The rig testing was supplemented with the numerical predictions. Results reveal that the throttle characteristic of DP indicates higher pressure rise and the loss reduction in tip clearance is mainly responsible for the performance improvement. For the compressor with DP, the pressure and flow angle are more uniform on exit plane. What’s more, the rotor with diffused passage reveals more robust than the conventional rotor at double clearance gap. Furthermore, the experimental data indicate that DP presents higher pressure rise at design and part speeds. At design speed, the stall margin was extended by 7.25%. Moreover, peak adiabatic efficiency of DP is also higher than that of CNV by about 0.7%.

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