Design and Development of a Two Stage Transonic Axial Flow Compressor

1996 ◽  
Author(s):  
Katsushi Nagai ◽  
Kazuaki Ikesawa ◽  
Takao Sugimoto ◽  
Toshinao Tanaka ◽  
Hiroshi Umino ◽  
...  
Keyword(s):  
Gas Turbines ◽  
Design Method ◽  
Axial Flow ◽  
Test Facility ◽  
Blade Design ◽  
Two Stage ◽  
Circular Arc ◽  
Axial Flow Compressor ◽  
Transonic Compressor ◽  
Flow Compressor

A highly loaded two stage transonic axial flow compressor, which forms a front stages of a multi stage compressor for industrial gas turbines, has been designed and tested. Overall pressure ratio is 2.25 and the first stage rotor tip Mach number is 1.15. Two airfoil types, Double Circular Arc airfoil and Multi Circular Arc airfoil, were designed for a transonic rotor blade under the same condition. MCA blade design method was devised and introduced. The blade design relied heavily on CFD techniques using a Euler code and a Navier Stokes code to cope with a precise treatment. The rig test was conducted by our compressor test facility to verify a validity of the transonic compressor design method and to compare the performance of the DCA and the MCA airfoils. This report describes the aerodynamic design and the test results as well as the test facility and instrumentation.

Inverse Design of Axial-Flow Compressor Blades Using Elastic Surface Algorithm in Subsonic and Transonic Flow Regimes With Separation

2016 ◽  
Author(s):  
M. H. Noorsalehi ◽  
M. Nili-Ahamadabadi ◽  
E. Shirani ◽  
M. Safari
Keyword(s):  
Pressure Distribution ◽  
Transonic Flow ◽  
Design Method ◽  
Axial Flow ◽  
Flow Regimes ◽  
Inverse Design ◽  
Elastic Surface ◽  
Axial Flow Compressor ◽  
Inverse Design Method ◽  
Flow Compressor

In this study, a new inverse design method called Elastic Surface Algorithm (ESA) is developed and enhanced for axial-flow compressor blade design in subsonic and transonic flow regimes with separation. ESA is a physically based iterative inverse design method that uses a 2D flow analysis code to estimate the pressure distribution on the solid structure, i.e. airfoil, and a 2D solid beam finite element code to calculate the deflections due to the difference between the calculated and target pressure distributions. In order to enhance the ESA, the wall shear stress distribution, besides pressure distribution, is applied to deflect the shape of the airfoil. The enhanced method is validated through the inverse design of the rotor blade of the first stage of an axial-flow compressor in transonic viscous flow regime. In addition, some design examples are presented to prove the effectiveness and robustness of the method. The results of this study show that the enhanced Elastic Surface Algorithm is an effective inverse design method in flow regimes with separation and normal shock.

Flow Studies on a Single Stage Transonic Axial Flow Compressor Retrofitted With Circumferential Grooves and Varied Rotor-Stator Axial Gap

2017 ◽  
Author(s):  
Anand P. Darji ◽  
Dilipkumar Bhanudasji Alone ◽  
Chetan S. Mistry
Keyword(s):  
Numerical Study ◽  
Computational Study ◽  
Axial Flow ◽  
Single Stage ◽  
Experimental Results ◽  
Stall Margin ◽  
Axial Flow Compressor ◽  
Transonic Compressor ◽  
Stall Margin Improvement ◽  
Flow Compressor

A transonic axial flow compressor undergoes severe vibrations due to instabilities like stall and surge when it operates at lower mass flow rate in the absence of any control devices. In present study, the attempt was made to understand the combine impact of circumferential casing grooves (CCG) of constant aspect ratio and different axial spacing between rotor and stator on the operating stability of single stage transonic axial compressor and that of rotor alone using numerical simulation. The optimum rotor-stator gap in the presence of grooved casing treatment was identified. The steady state numerical analysis was performed by using three-dimensional Reynolds Average Navier-Stokes equation adapting shear stress transport (SST) k-ω turbulence model. The study is reported in two sections. First section includes the detailed numerical study on baseline case having smooth casing wall (SCW). The computational results were validated with the experimental results available at Propulsion Division of CSIR-NAL, Bangalore. The computational study shows good agreement with experimental results. The second section comprises the effects of optimum designs of CCG and various axial spacing on the stall margin improvement of transonic compressor. Current computational study shows that the axial spacing between rotor and stator is an important parameter for improvement in stall margin not only for SCW but also for CCG. Therefore, the highest stall margin improvement of 9% has achieved for 75% axial spacing.

Experimental Investigation on the Effect of Porosity of Bend Skewed Casing Treatment on a Single Stage Transonic Axial Flow Compressor

2014 ◽  
Author(s):  
Dilipkumar B. Alone ◽  
Subramani Satish Kumar ◽  
Shobhavathy Thimmaiah ◽  
Janaki Rami Reddy Mudipalli ◽  
A. M. Pradeep ◽  
...  
Keyword(s):  
Experimental Investigation ◽  
Axial Flow ◽  
Compressor Stage ◽  
Casing Treatment ◽  
Operating Range ◽  
Axial Flow Compressor ◽  
Transonic Compressor ◽  
Stable Operating ◽  
Flow Compressor ◽  
Casing Treatments

A bend skewed casing treatment was designed, to study the influence of one of its geometrical parameter porosity on the stable performance of single stage transonic axial flow compressor. The compressor was designed for the stage total-to-total pressure ratio of 1.35, corrected mass flow rate of 22 kg/s at corrected design speed of 12930 RPM. Bend skewed casing treatment has an axial inlet segment till 50% of the total length and rear segment that is skewed by 45° in the direction of the rotor tip section stagger. Both the sections were oriented at a skew angle of 45° to the radial plane such that the flow exiting the slot is in counter-clockwise direction to that of the rotor direction. The casing treatment slot width was equal to the maximum thickness of the rotor blades. Three casing treatment configurations were identified for the current experimental investigation. All the treatment geometries considered for the experimental research have lower porosities than reported in the open literatures. The effect of the porosity parameter on the performance of transonic compressor stage was evaluated at two axial coverages of 20% and 40% relative to the rotor tip axial chord. Performance maps were obtained for the solid casing and casing treatment with three different porosities. Comparative studies were carried out and experimental results showed a maximum of 65% improvement in the stable operating range of the compressor for one of the treatment configurations. It was also observed that the stable operating range of the compressor increases with an increase in the casing treatment porosity. All the casing treatment configurations showed that the compressor stall occurs at lower mass flows as compared to the solid casing. Compressor stage peak efficiency shows significant degradations with increase in the porosity as compared to solid casing. Detailed blade element performances were also obtained using calibrated multi-hole aerodynamic probe. Comparative variations of flow parameters like absolute flow angle, Mach number were studied at full flow and near stall conditions for the solid casing and casing treatment configurations. Hot wire measurements show very high fluctuation in the inlet axial velocity in the presence of solid casing as compared to casing treatments. Experimental investigation revealed that the porosity of the casing treatments has strong influence on the transonic compressor stage performance.

Experimental investigations on instability evolution in a transonic compressor at different rotor speeds

2015 ◽  
Vol 229 (18) ◽  
pp. 3378-3391 ◽  
Author(s):  
Qiushi Li ◽  
Tianyu Pan ◽  
Tailu Sun ◽  
Zhiping Li ◽  
Yifang Gong
Keyword(s):  
Low Frequency ◽  
Axial Flow ◽  
Rotating Stall ◽  
Operating Conditions ◽  
Experimental Investigations ◽  
Rotor Speed ◽  
Axial Flow Compressor ◽  
Transonic Compressor ◽  
New Type ◽  
Flow Compressor

Experimental investigations are conducted to study the instability evolution in a transonic axial flow compressor at four specific rotor speeds covering both subsonic and transonic operating conditions. Two routes of evolution to final instability are observed in the test compressor: at low rotor speeds, a disturbance in the rotor tip region occurs and then leads to rotating stall, while at high rotor speeds, a low-frequency disturbance in the hub region leads the compressor into instability. Different from stall and surge, this new type of compressor instability at high rotor speed is initiated through the development of a low-frequency axisymmetric disturbance at the hub, and we name it “partial surge”. The frequency of this low-frequency disturbance is approximately the Helmholtz frequency of the system and remains constant during instability inception. Finally, a possible mechanism for the occurrence of different instability evolutions and the formation of partial surge are also discussed.

Experimental Investigation of Stepped Tip Gap Effects on the Performance of a Transonic Axial-Flow Compressor Rotor

1998 ◽  
Vol 120 (3) ◽  
pp. 477-486 ◽  
Author(s):  
D. W. Thompson ◽  
P. I. King ◽  
D. C. Rabe
Keyword(s):  
Mass Flow ◽  
Pressure Ratio ◽  
Axial Flow ◽  
Operating Conditions ◽  
Tip Clearance ◽  
Axial Flow Compressor ◽  
Transonic Compressor ◽  
Compressor Rotor ◽  
Percent Improvement ◽  
Flow Compressor

The effects of stepped-tip gaps and clearance levels on the performance of a transonic axial-flow compressor rotor were experimentally determined. A two-stage compressor with no inlet guide vanes was tested in a modern transonic compressor research facility. The first-stage rotor was unswept and was tested for an optimum tip clearance with variations in stepped gaps machined into the casing near the aft tip region of the rotor. Nine causing geometries were investigated consisting of three step profiles at each of three clearance levels. For small and intermediate clearances, stepped tip gaps were found to improve pressure ratio, efficiency, and flow range for most operating conditions. At 100 percent design rotor speed, stepped tip gaps produced a doubling of mass flow range with as much as a 2.0 percent increase in mass flow and a 1.5 percent improvement in efficiency. This study provides guidelines for engineers to improve compressor performance for an existing design by applying an optimum casing profile.

Experimental Investigation of Stepped Tip Gap Effects on the Performance of a Transonic Axial-Flow Compressor Rotor

1997 ◽  
Author(s):  
Donald W. Thompson ◽  
Paul I. King ◽  
Douglas C. Rabe
Keyword(s):  
Mass Flow ◽  
Pressure Ratio ◽  
Axial Flow ◽  
Operating Conditions ◽  
Tip Clearance ◽  
Axial Flow Compressor ◽  
Transonic Compressor ◽  
Compressor Rotor ◽  
Compressor Performance ◽  
Flow Compressor

The effects of stepped tip gaps and clearance levels on the performance of a transonic axial-flow compressor rotor were experimentally determined. A two-stage compressor with no inlet guide vanes was tested in a modern transonic compressor research facility. The first-stage rotor was unswept and was tested for an optimum tip clearance with variations in stepped gaps machined into the casing near the aft tip region of the rotor. Nine casing geometries were investigated consisting of three step profiles at each of three clearance levels. For small and intermediate clearances, stepped tip gaps were found to improve pressure ratio, efficiency, and flow range for most operating conditions. At 100% design rotor speed, stepped tip gaps produced a doubling of mass flow range with as much as a 2.0% increase in mass flow and a 1.5% improvement in efficiency. This study provides guidelines for engineers to improve compressor performance for an existing design by applying an optimum casing profile.

Large-Scale DES Analysis of Unsteady Flow Field in a Multi-Stage Axial Flow Compressor at Off-Design Condition Using K Computer

2015 ◽  
Author(s):  
Kazutoyo Yamada ◽  
Masato Furukawa ◽  
Satoshi Nakakido ◽  
Akinori Matsuoka ◽  
Kentaro Nakayama
Keyword(s):  
Boundary Condition ◽  
Unsteady Flow ◽  
Gas Turbines ◽  
Large Scale ◽  
Axial Flow ◽  
Axial Flow Compressor ◽  
Eddy Simulation ◽  
Multi Stage ◽  
Flow Phenomena ◽  
Flow Compressor

The paper presents the results of large-scale numerical simulations which were conducted for better understanding of unsteady flow phenomena in a multi-stage axial flow compressor at off-design condition. The compressor is a test rig compressor which was used for development of the industrial gas turbine, Kawasaki L30A. The compressor consists of 14 stages, the front two stages and the front half stages of which were investigated in the present study. The final goal of this study is to elucidate the flow mechanism of the rotating stall inception in the multi-stage axial compressor for actual gas turbines, and according to the test data it is considered that the 2nd stage and the 5th or 6th stage are suspected of leading to the stall. In order to capture precise flow physics in the compressor, a computational mesh for the simulation was generated to have at least several million cells per passage, which amounted to 650 million cells for the front 2-stage simulation and two billion cells for the front 7-stage simulation (about three hundred million cells for each stage). Since these were still not enough for the large-eddy simulation (LES), the detached-eddy simulation (DES) was employed, which can calculate flow fields except near-wall region by LES. The required computational resources were quite large for such simulations, so the computations were conducted on the K computer (RIKEN AICS in Japan). The simulations were well validated, showing good agreement with the measurement results obtained in the test. In the validation, the effect of the boundary condition for the casing wall was also investigated by comparing the results between the adiabatic boundary condition and the isothermal boundary condition. As for the unsteady effect, the wake/blade interaction was investigated in detail. In addition, unsteady flow phenomena in the present compressor at off-design condition were analyzed by using data mining techniques such as vortex identification and limiting streamline drawing with the LIC (line integral convolution) method. The simulation showed that they could be caused by the corner separation on the hub side.

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