3D CFD Compressor Map Computation of a Multi-Stage Axial Compressor With Off-Design Adjusted Rotor Geometries

2016 ◽  
Author(s):  
Christian Janke ◽  
Markus Goller ◽  
Ivo Martin ◽  
Lilia Gaun ◽  
Dieter Bestle
Keyword(s):  
High Speed ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Operating Conditions ◽  
Tip Clearance ◽  
Design Point ◽  
Multi Stage ◽  
Compressor Design ◽  
Compressor Map ◽  
Aero Engines

Compressor maps of aero engines show the relation between corrected mass flow, corrected shaft speed, pressure ratio, and efficiency, where different operating conditions of the compressor are represented by different speed lines. These speed lines are an important information for the compressor design process, since they show important operation bounds like surge and choke. Typically, 3D CFD compressor maps are computed with the so called hot geometry given by the aerodynamic design point. But in reality aerofoil shapes change depending on engine speeds and gas loads resulting in twist of the blades and changes of tip clearance. In order to obtain a higher quality compressor map, all these effects must be taken into account. Therefore, a process is utilized which uses coupled CFD and FE analyses to account for load adjusted geometries aside the design point. For transformation of FE results into the CFD model a cold-to-hot blade morphing technique is used. The studies are performed for a 4.5 stage high speed axial compressor, where effects of varying tip clearance and geometry deformation are considered separately from each other. Finally, their combined effects are studied.

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.

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

2012 ◽  
Vol 225 ◽  
pp. 233-238
Author(s):  
A.M. Pradeep ◽  
R.N. Chiranthan ◽  
Debarshi Dutta ◽  
Bhaskar Roy
Keyword(s):  
Rotor Blade ◽  
Cross Flow ◽  
Axial Compressor ◽  
Axial Flow ◽  
Pressure Rise ◽  
Tip Clearance ◽  
Suction Surface ◽  
Design Point ◽  
Compressor Rotor ◽  
Multi Stage

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.

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.

Effects of Tip Clearance and Operating Conditions on the Flow Structure and Turbulence Within an Axial Compressor Rotor Passage

2016 ◽  
Author(s):  
Yuanchao Li ◽  
Huang Chen ◽  
David Tan ◽  
Joseph Katz
Keyword(s):  
Flow Structure ◽  
High Speed ◽  
Pressure Coefficient ◽  
Axial Compressor ◽  
Suction Side ◽  
Secondary Structures ◽  
Operating Conditions ◽  
Rotor Blades ◽  
Tip Clearance ◽  
Wide Gap

The flows in the tip regions of two rotors with blades of similar geometry but different tip clearance are studied experimentally to determine the effect of gap on the flow structure at different operating conditions. The experiments have been performed in the JHU optically index-matched facility, where the refractive index of the fluid is matched with that of the acrylic rotor blades and casing, facilitating unobstructed Stereo Particle Image Velocimetry (SPIV) measurements. The blade geometries are based on the first one and a half stages of the Low Speed Axial Compressor (LSAC) facility at NASA Glenn. The tip gap sizes are 0.49% and 2.3% of the blade chordlength, and measurements are performed for two flow rates, the lower of which is just above stall conditions. The presence and trajectories of the tip leakage vortex (TLV) and secondary structures are visualized by recording high speed movies of cavitation at lower pressures. The results consist of performance curves, distributions of velocity, circumferential vorticity and turbulent kinetic energy, as well as the strength and trajectory of vortices. Increasing the tip gap reduces the static-to-static pressure coefficient for all flow conditions. For the higher flow rate, a wider tip gap has several effects: (i) It delays the rollup of the TLV and its detachment from the suction side (SS) corner of the blade, presumably due to the larger distance from the endwall casing and the ‘image vortex’. (ii) It alters the blade loading and reduces the circulation shed from the blade. (iii) It delays the onset of TLV bursting in the aft part of the rotor passage. (iv) For both gaps, the endwall boundary layer separates at the point where the leakage flow meets the opposite-direction main passage flow. For the wide gap, the separated layer with opposite sign vorticity remains above the TLV; while for the narrow gap, the TLV entrains this layer around itself. And (v) consistent with the major differences in flow structure, the spatial distributions and magnitudes of all the turbulence intensity are also very different. Trends and flow structure are quite different at pre-stall conditions. Most notably, TLV rollup is still delayed for the wide gap, but vortex bursting and associated arrival of multiple secondary structures to the pressure side (PS) of the next blade occur earlier. Consequently, the turbulence level on both sides of the blade tip is substantially higher, and remnants of the previous TLV are ingested into the next tip gap.

Viscous Throughflow Modelling for Multi-Stage Compressor Design

1992 ◽  
Author(s):  
M. A. Howard ◽  
S. J. Gallimore
Keyword(s):  
Shear Force ◽  
High Speed ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Design System ◽  
Flow Angle ◽  
Axial Compressors ◽  
Multi Stage ◽  
Viscous Shear ◽  
Compressor Design

An existing throughflow method for axial compressors, which accounts for the effects of spanwise mixing using a turbulent diffusion model, has been extended to include the viscous shear force on the endwall. The use of a shear force, consistent with a no-slip condition, on the annulus walls in the throughflow calculations allows realistic predictions of the velocity and flow angle profiles near the endwalls. The annulus wall boundary layers are therefore incorporated directly in the throughflow prediction. This eliminates the need for empirical blockage factors or independent annulus boundary layer calculations. The axisymmetric prediction can be further refined by specifying realistic spanwise variations of loss coefficient and deviation to model the three-dimensional endwall effects. The resulting throughflow calculation gives realistic predictions of flow properties across the whole span of a compressor. This is confirmed by comparison with measured data from both low and high speed multi-stage machines. The viscous throughflow method has been incorporated into an axial compressor design system. The method predicts the meridional velocity defects in the endwall region and consequently blading can be designed which allows for the increased incidence, and low dynamic head, near to the annulus walls.

Performance Desensitization for a High-Speed Axial Compressor

2018 ◽  
Author(s):  
Yassine Souleimani ◽  
Huu Duc Vo ◽  
Hong Yu
Keyword(s):  
High Speed ◽  
Numerical Study ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Tip Clearance ◽  
Lower Sensitivity ◽  
Tip Leakage Flow ◽  
Stall Margin ◽  
Casing Treatment ◽  
Compressor Rotor

The increase in compressor tip clearance over the lifespan of an aero-engine leads to a long-term degradation in its fuel consumption and operating envelope. A highly promising recent numerical study on a theoretical high-speed axial compressor rotor proposed a novel casing treatment to decrease performance and stall margin sensitivity to tip clearance increase. This paper aims to apply and analyze, through CFD simulations, this casing treatment concept to a representative production axial compressor rotor with inherently lower sensitivity to tip clearance increase and complement the explanation on the mechanism behind the reduction in sensitivity. Simulations of the baseline rotor showed that the lower span region contribute as much to the pressure ratio sensitivity as the tip region which is dominated by tip leakage flow. In contrast, the efficiency sensitivity is mainly driven by losses occurring in the tip region. The novel casing treatment was successfully applied to the baseline rotor through a design refinement. Although the casing treatment causes some penalty in nominal performance, it completely reversed the pressure ratio sensitivity (i.e. pressure ratio increases with tip clearance) and reduced the efficiency sensitivity. The reversed pressure ratio sensitivity is explained by a rotation in the core flow in the lower span region indirectly induced by the flow injection from the casing treatment. The lower efficiency sensitivity comes from a reduction in the amount of fluid that crosses the tip clearance of two adjacent blades, known as double leakage. The casing treatment’s beneficial effect on stall margin sensitivity is less obvious because of the stall inception type of the baseline rotor and its change in the presence of the casing treatment.

3D CFD Compressor Map Computation Process Accounting for Geometry Changes due to Off-Design Operation Loads

2014 ◽  
Author(s):  
Christian Janke ◽  
Kai Karger ◽  
Lilia Gaun ◽  
Dieter Bestle ◽  
André Huppertz
Keyword(s):  
Design Process ◽  
Total Pressure ◽  
Pressure Ratio ◽  
Operating Conditions ◽  
Aerodynamic Design ◽  
Quality Performance ◽  
Computation Process ◽  
Compressor Map ◽  
Aero Engines ◽  
Total Pressure Ratio

Compressor maps of aero engines show the relation between corrected inlet mass flow and total pressure ratio for various engine speeds. Different speed lines represent different operating conditions of the compressor, where especially operating bounds like surge and choke are important for the design process. Typically, 3D CFD compressor maps are computed with the so called hot geometry given for the aerodynamic design point. However, in reality airfoil shapes will change for different engine speeds and gas loads resulting in twisted airfoils and changed tip clearances. Thus, using the nominal hot geometry for the whole compressor map is not fully correct. In order to obtain higher quality performance maps these effects need to be considered. The paper shows a process for computing compressor maps with 3D CFD, where strucural deformations of the blade due to varying speeds and gas loads are taken into account by blade morphing. This process is applied to a 1.5-stage compressor showcase.

Viscous Throughflow Modeling of Axial Compressor Bladerows Using a Tangential Blade Force Hypothesis

1998 ◽  
Vol 120 (4) ◽  
pp. 662-670 ◽  
Author(s):  
S. J. Gallimore
Keyword(s):  
High Speed ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Operating Conditions ◽  
Tip Clearance ◽  
Boundary Layer Flows ◽  
Quick Method ◽  
Novel Approach ◽  
Performance Predictions ◽  
Additional Losses

This paper describes the modeling of axial compressor blade rows in an axisymmetric viscous throughflow method. The basic method, which has been reported previously, includes the effects of spanwise mixing, using a turbulent diffusion model, and endwall shear within the throughflow calculation. The blades are modeled using a combination of existing two-dimensional blade performance predictions for loss and deviation away from the annulus walls and a novel approach using tangential blade forces in the endwall regions. Relatively simple assumptions about the behavior of the tangential static pressure force imposed by the blades allow the secondary deviations produced by tip clearance flows and the boundary layer flows at fixed blade ends to be calculated in the axisymmetric model. Additional losses are assigned in these regions based on the calculated deviations. The resulting method gives realistic radial distributions of loss and deviation across the whole span at both design and off-design operating conditions, providing a quick method of estimating the magnitudes of these effects in the preliminary design process. Results from the method are compared to measured data in low and high-speed compressors and multistage three-dimensional viscous CFD predictions.

Experimental and Numerical Investigation of Operating Range Enhancement Techniques in Centrifugal Compressor for Turbochargers

2017 ◽  
Author(s):  
Kishore Kumar Chandramohan ◽  
Kirubakaran Purushothaman ◽  
Vidyadheesh Pandurangi ◽  
Kishore Prasad Deshkulkarni
Keyword(s):  
High Speed ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Centrifugal Compressors ◽  
Operating Range ◽  
Auxiliary Power Units ◽  
Flow Mechanisms ◽  
Cruise Missiles ◽  
Compressor Map ◽  
Ported Shroud

High speed centrifugal compressors are used in turbochargers and in small gas turbine engines that typically power cruise missiles, helicopters and auxiliary power units (APU). Centrifugal compressors have wider operating range compared to axial compressor and are compact. Though centrifugal compressors having a pressure ratio of the order of 12:1 per stage have been demonstrated with reasonably good isentropic efficiencies, achieving a wider operating range has always been a challenge. A Turbocharger that needs to be designed to function both at sea-level and 5 km altitude conditions, requires a wider compressor map to accommodate the diesel engine operating line. A wider compressor map can be achieved by various techniques. The approaches used in the present study include providing pinch in the diffuser entry region and ported shroud arrangement in the compressor casing. A parametric study has been carried out by varying geometric parameters and an appropriate configuration that offers lower total pressure loss and better diffuser pressure recovery is chosen. The flow mechanisms responsible for better performance is investigated numerically for various configurations with diffuser pinch. To further enhance the operating range, a ported shroud configuration in the compressor housing is designed and analysed with the finalized diffuser pinch. Results of computational analysis for different ported shroud slot geometries have been studied numerically and are presented. Two configurations have been tested in a turbo-drive based test rig. The first configuration is only with diffuser pinch and the second configuration is with diffuser pinch and ported shroud. The extent of map width enhancement obtained by each technique is presented and compared with numerical analysis. The test results show good match with the predicted trend and confirms that diffuser pinch and ported shroud configurations offer significant enhancement in achieving a wider operating range. The flow mechanisms responsible are discussed in detail in the paper.

Viscous Throughflow Modelling of Axial Compressor Bladerows Using a Tangential Blade Force Hypothesis

1997 ◽  
Author(s):  
Simon J. Gallimore
Keyword(s):  
High Speed ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Operating Conditions ◽  
Tip Clearance ◽  
Boundary Layer Flows ◽  
Quick Method ◽  
Novel Approach ◽  
Performance Predictions ◽  
Additional Losses

This paper describes the modelling of axial compressor blade rows in an axisymmetric viscous throughflow method. The basic method, which has been reported before, includes the effects of spanwise mixing, using a turbulent diffusion model, and endwall shear within the throughflow calculation. The blades are modelled using a combination of existing two-dimensional blade performance predictions for loss and deviation away from the annulus walls and a novel approach using tangential blade forces in the endwall regions. Relatively simple assumptions about the behaviour of the tangential static pressure force imposed by the blades allow the secondary deviations produced by tip clearance flows and the boundary layer flows at fixed blade ends to be calculated in the axisymmetric model. Additional losses are assigned in these regions based on the calculated deviations. The resulting method gives realistic radial distributions of loss and deviation across the whole span at both design and off-design operating conditions providing a quick method of estimating the magnitudes of these effects in the preliminary design process. Results from the method are compared to measured data in low and high speed compressors and multistage three-dimensional viscous CFD predictions.

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