Time Resolved Simulation and Experimental Validation of the Flow in Axial Slot Casing Treatments for Transonic Axial Compressors

2008 ◽  
Author(s):  
Giovanni A. Brignole ◽  
Florian C. T. Danner ◽  
Hans-Peter Kau
Keyword(s):  
Mass Flow ◽  
Total Pressure ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Compressor Stage ◽  
Time Resolved ◽  
Transonic Compressor ◽  
Design Speed ◽  
Casing Treatments ◽  
Total Pressure Ratio

Building on the experience of previous investigations, a casing treatment was developed and applied to an axial transonic compressor stage, in literature referred to as Darmstadt Rotor 1. The aerodynamics of the experimental compressor stage was improved by applying axially orientated semicircular slots to the original plain casing, which both enhanced the operating range and design point efficiency. A gain in total pressure ratio along the entire design speed line was also observed. Within the scope of this study four different axial casing treatments were designed. Their effect on the flow in a transonic axial compressor stage was investigated parametrically using time-resolved 3D-FANS simulations with a mesh of approximately 4.8 · 106 grid points. This research aims to identify correlations between the geometrical cavity design and the changed channel flow. The findings help to formulate parameters for evaluating the performance of casing treatments. These criteria can further be used as target functions in the design optimisation process. The predicted behaviour of the transonic compressor was validated against experiments as well as an alternative numerical model, the non-linear harmonic method. Both confirmed the effect of the slots in raising efficiency as well as moving the design speed line towards higher pressure ratios. In the experiments, the addition of the slots increased the total pressure ratio at stall conditions by more than 5% and reduced mass flow from 87.5% of the design mass flow to less than 77.5% compared to the original geometry.

Experimental and Numerical Analysis of Axial Skewed Slot Casing Treatments for a Transonic Compressor Stage

2009 ◽  
Author(s):  
Florian C. T. Danner ◽  
Hans-Peter Kau ◽  
Martin M. Mu¨ller ◽  
Heinz-Peter Schiffer ◽  
Giovanni A. Brignole
Keyword(s):  
Total Pressure ◽  
Pressure Ratio ◽  
Computational Domain ◽  
Compressor Stage ◽  
Time Resolved ◽  
Transonic Compressor ◽  
Computational Fluid Dynamics Analysis ◽  
Grid Points ◽  
Casing Treatments ◽  
Total Pressure Ratio

An investigation of a single-stage transonic compressor with axial skewed slot casing treatments is presented. The studied compressor stage is characterised by a design mass flow rate of 16 kg/s at a total pressure ratio of 1.5 and a rotor tip speed of 400m/s. The research comprises experimental measurements as well as time-resolved simulations at full and part speed. Total pressure ratio measurements and efficiency speedlines are complemented by traversing downstream of the stator and static pressure measurements at the rotor end wall. The experimental work is supported by unsteady computational fluid dynamics analysis to provide further insight into the ruling flow phenomena. The simulations were carried out fully three-dimensionally in a computational domain with approximately 4.8 million grid points including the cavity mesh. The application of the axial skewed slots led to both, an enhanced operating range and an increased design point efficiency. Rises in total pressure ratio along the entire speed lines were observed.

Numerical Analysis to Investigate the Effect of Swept Rotor on the Overall Performance of a Transonic Compressor Stage

2014 ◽  
Author(s):  
Shobhavathy M. Thimmaiah ◽  
Ramesha Gurikelu ◽  
Nisha Sherief
Keyword(s):  
Flow Field ◽  
Total Pressure ◽  
Pressure Ratio ◽  
Stability Margin ◽  
Compressor Stage ◽  
Operating Range ◽  
Transonic Compressor ◽  
Overall Performance ◽  
Design Speed ◽  
The Stability

This paper presents the steady state numerical analyses carried out to investigate the effect of forward and backward swept rotor on the overall performance and stability margin of single stage transonic axial flow compressor. Initially, the analyses were carried out on a radially stacked rotor/baseline configuration and obtained the overall performance map of the compressor stage. These results were compared with the available experimental data for validation. Further, investigations were carried out on geometrically modified rotor with six configurations having 5, 10 and 15° forward and backward sweep. A commercial 3-Dimensional CFD package, ANSYS FLUENT was used to compute the complex flow field of transonic compressor rotors. The flow field structures were studied with the help of Mach number total pressure contours. The results of modified rotor geometry indicated that the peak adiabatic efficiency and the total pressure ratio for all the tested forward and backward swept rotor configurations are marginally higher than that of the baseline configuration at all speeds. The operating ranges of all the swept rotor configurations are found to be higher than that of the baseline configuration. The operating range is broader at lower operating speeds than at design speed condition. Rotor with 10° forward sweep and 5° backward sweep indicated the noteworthy improvement in the operating range against the baseline configuration. The stability margin of 11.3, 6.6, 5.2 and 3.5% at 60, 80, 90 and 100% of the design speed respectively compared to the baseline configuration obtained from 10° forward sweep. Rotor with 5° backward sweep showed the stability margin of 12, 4, 3.9 and 3% at 60, 80, 90 and 100% of the design speed respectively compared to the baseline configuration.

Aerodynamic Design and Analysis of a High Pressure Ratio Aspirated Compressor Stage

2000 ◽  
Author(s):  
Ali A. Merchant ◽  
Mark Drela ◽  
Jack L. Kerrebrock ◽  
John J. Adamczyk ◽  
Mark Celestina
Keyword(s):  
Boundary Layer ◽  
Total Pressure ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Adverse Pressure Gradient ◽  
Compressor Stage ◽  
High Pressure Ratio ◽  
Total Pressure Ratio ◽  
Isentropic Efficiency

The pressure ratio of axial compressor stages can be significantly increased by controlling the development of blade and endwall boundary layers in regions of adverse pressure gradient by means of boundary layer suction. This concept is validated and demonstrated through the design and analysis of a unique aspirated compressor stage which achieves a total pressure ratio of 3.5 at a tip speed of 1500 ft/s. The aspirated stage was designed using an axisymmetric through-flow code coupled with a quasi three-dimensional cascade plane code with inverse design capability. Validation of the completed design was carried out with three-dimensional Navier-Stokes calculations. Spanwise slots were used on the rotor and stator suction surfaces to bleed the boundary layer with a total suction requirement of 4% of the inlet mass flow. Additional bleed of 3% was also required on the hub and shroud near shock impingement locations. A three-dimensional viscous evaluation of the design showed good agreement with the quasi three-dimensional design intent, except in the endwall regions. The three-dimensional viscous analysis predicted a mass averaged total pressure ratio of 3.7 at an isentropic efficiency of 93% for the rotor, and a mass averaged total pressure ratio of 3.4 at an isentropic efficiency of 86% for the stage.

Effect of Backward Sweep on Aerodynamic Performance of a 1.5-Stage Highly Loaded Axial Compressor

2020 ◽  
Author(s):  
Song Huang ◽  
Chuangxin Zhou ◽  
Chengwu Yang ◽  
Shengfeng Zhao ◽  
Mingyang Wang ◽  
...  
Keyword(s):  
Total Pressure ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Aerodynamic Performance ◽  
Blade Loading ◽  
Stall Margin ◽  
Axial Compressors ◽  
Design Speed ◽  
Total Pressure Ratio ◽  
Highly Loaded

Abstract As a degree of freedom in the three-dimensional blade design of axial compressors, the sweep technique significantly affects the aerodynamic performance of axial compressors. In this paper, the effects of backward sweep rotor configurations on the aerodynamic performance of a 1.5-stage highly loaded axial compressor at different rotational design speeds are studied by numerical simulation. The aim of this work is to improve understanding of the flow mechanism of backward sweep on the aerodynamic performance of a highly loaded axial compressor. A commercial CFD package is employed for flow simulations and analysis. The study found that at the design rotational speed, compared with baseline, backward sweep rotor configurations reduce the blade loading near the leading edge but slightly increases the blade loading near the trailing edge in the hub region. As the degree of backward sweep increases, the stall margin of the 1.5-stage axial compressor increase first and then decrease. Among different backward sweep rotor configurations, the 10% backward sweep rotor configuration has the highest stall margin, which is about 2.5% higher than that of baseline. This is due to the change of downstream stator incidence, which improves flow capacity near the hub region. At 80% rotational design speed, backward sweep rotor configurations improve stall margin and total pressure ratio of the compressor. It’s mainly due to the decreases of the rotor incidence near the middle span, which results in the decreases of separation on the suction surface. At 60% rotational design speed, detached shock disappears. Backward sweep rotor configurations deteriorate stall margin of the compressor, but increase total pressure ratio and adiabatic efficiency when the flow rate is lower than that at peak efficiency condition. Therefore, it’s necessary to consider the flow field structure of axial compressors at whole operating conditions in the design process and use the design freedom of sweep to improve the aerodynamic performance.

Effect of swirl on the performance and stability of transonic axial compressor

2017 ◽  
Vol 232 (6) ◽  
pp. 608-621
Author(s):  
Baofeng Tu ◽  
Luyao Zhang ◽  
Jun Hu
Keyword(s):  
Mass Flow ◽  
Total Pressure ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Maximum Efficiency ◽  
Stable Boundary ◽  
Tip Leakage ◽  
Stable Operating ◽  
Swirl Intensity ◽  
Total Pressure Ratio

To investigate the effect of twin swirl and bulk swirl on the performance and stability of a transonic axial compressor, a blade swirl generator was designed and simulated with a transonic single rotor using steady and unsteady numerical calculation methods. The bulk swirl intensity was adjusted by replacing the blades with different camber angles. The twin swirl intensity was decreased by reducing the blade number. The counter-rotating bulk swirl generated a significant drop in both the efficiency and stall margin, and resulted in an increase in the choked mass flow, and total pressure ratio. The co-rotating bulk swirl generated a decrease in the mass flow, total pressure ratio and stable operating range. The counter-rotating bulk swirl resulted in suction surface boundary layer separation beyond 50% of the span-wise height as well as more serious tip leakage blockage. The twin swirl resulted in a decrease in the total pressure ratio, maximum efficiency and stable operating range. The steady and unsteady numerical calculation results were consistent, though some differences were observed in the values. For the steady calculation, the maximum efficiency and choked mass flow decreased by 0.88% and 1.74%, respectively, and the mass flow at the stable boundary increased by 3.92% as compared to the uniform inlet flow at twin swirl intensity of 24°. During the unsteady calculation, the mass flow exhibited an increase of only 2.2% at the stable boundary. Under twin swirl and co-rotating bulk swirl and uniform inlet flow, the leading edge spillage of the tip leakage flow resulted in compressor instability. The counter-rotating bulk swirl changed the mechanism of instability. The characterisation of the swirl distortion presented a difference between the steady and unsteady calculations near the stable boundary. The unsteady calculation exhibited a lower mass flow at the stable boundary point and a higher total pressure ratio.

Research on Optimization of the Design(Inverse) Issue of S2 Surface of Axial Compressor Based on Genetic Algorithm

2020 ◽  
Author(s):  
Zijing Chen ◽  
Bo Liu ◽  
Xiaoxiong Wu
Keyword(s):  
Genetic Algorithm ◽  
Total Pressure ◽  
Fitness Function ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Local Data ◽  
Real Coded Genetic Algorithm ◽  
Overall Efficiency ◽  
Total Pressure Ratio ◽  
Peak Value

Abstract In order to further improve the effectiveness of design(inverse) issue of S2 surface of axial compressor, a design method of optimization model based on real-coded genetic algorithm is instructed, with a detailed description of some important points such as the population setting, the fitness function design and the implementation of genetic operator. The method mainly takes the pressure ratio, the circulation as the optimization variables, the total pressure ratio and the overall efficiency of the compressor as the constraint condition and the decreasing of the diffusion factor of the compressor as the optimization target. In addition, for the propose of controlling the peak value of some local data after the optimization, a local optimization strategy is proposed to make the method achieve better results. In the optimization, the streamline curvature method is used to perform the iterative calculation of the aerodynamic parameters of the S2 flow surface, and the polynomial fitting method is used to optimize the dimensionality of the variables. The optimization result of a type of ten-stage axial compressor shows that the pressure ratio and circulation parameters have significant effect on the diffusion factor’s distribution, especially for the rotor pressure ratio. Through the optimization, the smoothness of the mass-average pressure ratio distribution curve of the rotors at all stages of the compressor is improved. The maximum diffusion factors in spanwise of rotor rows at the first, fifth and tenth stage of the compressor are reduced by 1.46%, 12.53% and 8.67%, respectively. Excluding the two calculation points at the root and tip of the blade because of the peak value, the average diffusion factors in spanwise are reduced by 1.28%, 3.46%, and 1.50%, respectively. For the two main constraints, the changes of the total pressure ratio and overall efficiency are less than 0.03% and 0.032%, respectively. In the end, a 3-d CFD numerical result is given to testify the effects of the optimization, which shows that the loss in the compressor is decreased by the optimization algorithm.

Some Experiments With a Supersonic Axial Compressor Stage

1987 ◽  
Vol 109 (3) ◽  
pp. 388-397 ◽  
Author(s):  
A. J. Wennerstrom
Keyword(s):  
Pressure Ratio ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Design Flow ◽  
Vortex Generators ◽  
Boundary Layer Control ◽  
Compressor Stage ◽  
Design Speed ◽  
Isentropic Efficiency ◽  
Layer Control

Between 1970 and 1974, ten variants of a supersonic axial compressor stage were designed and tested. These included two rotor configurations, three rotor tip clearances, addition of boundary-layer control consisting of vortex generators on both the outer casing and the rotor, and the introduction of slots in the stator vanes. Design performance objectives were a stage total pressure ratio of 3.0 with an isentropic efficiency of 0.82 at a tip speed of 1600 ft/s (488 m/s). The first configuration passed only 70 percent of design flow at design speed, achieving a stage pressure ratio of 2.25 at a peak stage isentropic efficiency of 0.61. The rotor was grossly separated. The tenth variant passed 91.4 percent of design flow at design speed, producing a stage pressure ratio of 3.03 with an isentropic efficiency of 0.75. The rotor achieved a pressure ratio of 3.59 at an efficiency of 0.87 under the same conditions. Major conclusions were that design tools available today would undoubtedly permit the original goals to be met or exceeded. However, the application for such a design is currently questionable because efficiency goals considered acceptable for most current programs have risen considerably from the level considered acceptable at the inception of this effort. Splitter vanes placed in the rotor permitted very high diffusion levels to be achieved without stalling. However, viscous effects causing three-dimensional flows violating the assumption of flow confined to concentric stream tubes were so strong that a geometry optimization does not appear practical without a three-dimensional, viscous analysis. Passive boundary-layer control in the form of vortex generators and slots does appear to offer some benefit under certain circumstances.

Optimization of Highly Loaded Fan Rotor Based on Throughflow Model

2007 ◽  
Author(s):  
Hong Wu ◽  
Qiushi Li ◽  
Sheng Zhou
Keyword(s):  
Mass Flow ◽  
Total Pressure ◽  
Pressure Ratio ◽  
Optimization Method ◽  
Design Variables ◽  
Adaptive Simulated Annealing ◽  
Throughflow Model ◽  
Important Design ◽  
Total Pressure Ratio ◽  
Highly Loaded

This paper presents an optimization method for fan/compressor which couples throughflow model solving axisymmetric Euler equations with adaptive simulated annealing (ASA) algorithm. One of the advantages of this optimization method is that it spends much less time than 3D optimization due to the rapid solving of throughflow model. In addition, the optimization space is quite extensive because more design variables can be adjusted in throughflow phase, such as swirl distribution, hub curve and sweep. To validate this optimization method, a highly loaded fan rotor with pressure ratio of 3.06 as a baseline is optimized. During the optimization process, the objective function is total pressure ratio, moreover, mass flow and efficiency are selected as the constraint conditions. Three important design variables including swirl distribution, hub curve and sweep are parameterized using Bezier curve, and then optimized in throughflow model independently, finally the optimum designs are validated using 3D viscous CFD solver. It is shown that pressure ratio and rotor loading can be improved further through optimizing swirl distribution, however, hub and sweep curves take more effects on mass flow and efficiency respectively. The optimization results demonstrate the advantage and feasibility of this optimization method.

Design and Test of a Transonic Axial Splittered Rotor

2015 ◽  
Author(s):  
Garth V. Hobson ◽  
Anthony J. Gannon ◽  
Scott Drayton
Keyword(s):  
Mass Flow ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Design Procedure ◽  
Tip Clearance ◽  
Turbine Engine ◽  
Transonic Compressor ◽  
Compressor Rotor ◽  
Transonic Axial Compressor ◽  
Commercial Off The Shelf

A new design procedure was developed that uses commercial-off-the-shelf software (MATLAB, SolidWorks, and ANSYS-CFX) for the geometric rendering and analysis of a transonic axial compressor rotor with splitter blades. Predictive numerical simulations were conducted and experimental data were collected in a Transonic Compressor Rig. This study advanced the understanding of splitter blade geometry, placement, and performance benefits. In particular, it was determined that moving the splitter blade forward in the passage between the main blades, which was a departure from the trends demonstrated in the few available previous transonic axial compressor splitter blade studies, increased the mass flow range with no loss in overall performance. With a large 0.91 mm (0.036 in) tip clearance, to preserve the integrity of the rotor, the experimentally measured peak total-to-total pressure ratio was 1.69 and the peak total-to-total isentropic efficiency was 72 percent at 100 percent design speed. Additionally, a higher than predicted 7.5 percent mass flow rate range was experimentally measured, which would make for easier engine control if this concept were to be included in an actual gas turbine engine.

Numerical Investigation on the Effects of Circumferential Coverage of Injection in a Transonic Compressor With Discrete Tip Injection

2014 ◽  
Author(s):  
Wei Wang ◽  
Wuli Chu ◽  
Haoguang Zhang ◽  
Yanhui Wu
Keyword(s):  
Total Pressure ◽  
Pressure Ratio ◽  
Time Lag ◽  
Three Dimensional ◽  
Stability Margin ◽  
Transonic Compressor ◽  
Blade Tip ◽  
Tip Injection ◽  
The Stability ◽  
Total Pressure Ratio

Discrete tip injection upstream of the rotor tip is an effective technique to extend stability margin for a compressor system in an aeroengine. The current study investigates the effects of injectors’ circumferential coverage on compressor performance and stability using time-accurate three-dimensional numerical simulations for multi passages in a transonic compressor. The percentage of circumferential coverage for all the six injectors ranges from 6% to 87% for the five investigated configurations. Results indicate that circumferential coverage of tip injection can greatly affect compressor stability and total pressure ratio, but has little influence on adiabatic efficiency. The improvement of compressor total pressure ratio is linearly related with the increasing circumferential coverage. The unsteady flow fields show that there exists a non-ignorable time lag of the injection effects between the passage inlet and outlet, and blade tip loading will not decline until the injected flow reaches the passage outlet. Stability improves sharply with the increasing circumferential coverage when the coverage is less than 27%, but increases flatly for the rest. It is proven that the injection efficiency which is a measurement of averaged blockage decrement in the injected region is an effective guideline to predict the stability improvement.

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