A Methodology for Variable Geometry Optimization of Multistage Axial Compressors

2017 ◽  
Author(s):  
M. Eric Lyall ◽  
Fred J. Eisert ◽  
Douglas C. Rabe ◽  
Patrick M. Fleisher
Keyword(s):  
Design Of Experiments ◽  
Pressure Ratio ◽  
Geometry Optimization ◽  
Axial Compressor ◽  
Optimization Method ◽  
Variable Geometry ◽  
Factorial Design Of Experiments ◽  
Stage Matching ◽  
The Impact ◽  
Corrected Flow

This paper presents a procedure for experimentally optimizing a multistage axial compressor. Due to the usual proprietary nature of such tests, a mean-line model of a nine-stage compressor with three rows of variable geometry is used instead of a real machine as a testbed for explaining the optimization method. The compressor is optimized to achieve design-intent corrected flow and pressure ratio while achieving acceptable efficiency and stage matching. The optimization is performed using a response surface methodology that leverages a full factorial design of experiments approach. The resulting empirical models of compressor performance are of high quality, with coefficients of determination exceeding 0.99. An important finding of the work is that stage interactions are important for modeling both efficiency and stage matching, much more than for corrected flow and pressure ratio. Additionally the empirical equations resulting from the design of experiments analysis provide sensitivities due to changes in the variable geometry. These sensitivities can be applied to understanding the impact of uncertainties related to rigging the variable geometry and for assessing potential new or upgraded compressor designs.

scholarly journals Research on the Application of Partial Similarity for a 1-1/2 Axial Compressor

Processes ◽  
2020 ◽  
Vol 8 (9) ◽  
pp. 1121
Author(s):  
Hong Xie ◽  
Moru Song ◽  
Bo Yang
Keyword(s):  
Low Cost ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Optimization Method ◽  
Casing Treatment ◽  
Partial Similarity ◽  
Similarity Principle ◽  
Performance Curves ◽  
Flow Similarity ◽  
The Impact

In this paper, a method based on the partial similarity principle is presented to improve the aerodynamic design with low cost and high accuracy for a 1-1/2 axial compressor. By means of this method, during the process of a similar design, the machine Mach number and flowrate coefficient are maintained. The flow similarity between the prototype and its large-scaled alternative was observed, according to a detailed analysis of flow fields of rotor and stator. As well, the relative discrepancies of isentropic efficiency and pressure ratio between two models are 1.25% and 0.4% at design point, respectively. Besides, their performance curves agreed very well in the whole operating range. Moreover, it was also found that the flow similarity between the two models can be maintained under unsteady working conditions. Thereafter, in order to investigate the impact of stability optimization method on the similarity principle, casing treatment with single circumferential groove was applied to these two models. The flow similarity was still maintained and the flowrate near the stall was reduced about 1.1% with negligible deterioration of the overall performance.

The Impact of Manufacturing Variations on Performance of a Transonic Axial Compressor Rotor

2018 ◽  
Author(s):  
Marcus Lejon ◽  
Niklas Andersson ◽  
Lars Ellbrant ◽  
Hans Mårtensson
Keyword(s):  
Flow Field ◽  
Geometric Mean ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Mean Deviation ◽  
Design Intent ◽  
Compressor Rotor ◽  
Measuring Machine ◽  
The Mean ◽  
The Impact

In this paper, the impact of manufacturing variations on performance of an axial compressor rotor are evaluated at design rotational speed. The geometric variations from the design intent were obtained from an optical coordinate measuring machine and used to evaluate the impact of manufacturing variations on performance and the flow field in the rotor. The complete blisk is simulated using 3D CFD calculations, allowing for a detailed analysis of the impact of geometric variations on the flow. It is shown that the mean shift of the geometry from the design intent is responsible for the majority of the change in performance in terms of mass flow and total pressure ratio for this specific blisk. In terms of polytropic efficiency, the measured geometric scatter is shown to have a higher influence than the geometric mean deviation. The geometric scatter around the mean is shown to impact the pressure distribution along the leading edge and the shock position. Furthermore, a blisk is analyzed with one blade deviating substantially from the design intent, denoted as blade 0. It is shown that the impact of blade 0 on the flow is largely limited to the blade passages that it is directly a part of. The results presented in this paper also show that the impact of this blade on the flow field can be represented by a simulation including 3 blade passages. In terms of loss, using 5 blade passages is shown to give a close estimate for the relative change in loss for blade 0 and neighboring blades.

Simple Recuperated s-CO2 Cycle Revisited: Optimization of Operating Parameters for Maximum Cycle Efficiency

2019 ◽  
Author(s):  
Anchit Dutta ◽  
Adhip Gupta ◽  
Sharath Sathish ◽  
Aman Bandooni ◽  
Pramod Kumar
Keyword(s):  
Design Of Experiments ◽  
Design Space ◽  
Pressure Ratio ◽  
Cycle Performance ◽  
Maximum Temperature ◽  
Brayton Cycle ◽  
High Side ◽  
Side Pressure ◽  
Cycle Efficiency ◽  
The Impact

Abstract The paper presents modeling and Design of Experiments (DOE) analysis for a simple recuperated s-CO2 closed loop Brayton cycle operating at a maximum temperature of 600°C and a compressor inlet temperature of 45°C. The analysis highlights the impact of isentropic efficiencies of the turbine and compressor, decoupled in this case, on other equipment such as recuperator, gas cooler and heater, all of which have a bearing on the overall performance of the s-CO2 Brayton cycle. A MATLAB program coupled with REFPROP is used to perform the thermodynamic analysis of the cycle. A design space exploration with a Design of Experiments (DOE) study is undertaken using I-sight™ (multi-objective optimization software), which is coupled with the MATLAB code. The outcome of the DOE study provides the optimal pressure ratios and high side pressures for maximum cycle efficiency in the design space. By varying pressure ratios along with a floating high side pressure, the analysis reveals that the cycle performance exhibits a peak around a pressure ratio of 2.5, with cycle efficiency being the objective function. A further interesting outcome of the DOE study reveals that the isentropic efficiencies of the compressor and turbine have a strong influence not only on the overall cycle efficiency, but also the optimum pressure ratio as well as the threshold pressures (low as well as high side pressure). An important outcome of this exercise shows that the isentropic efficiency of the turbine has a much greater impact on the overall cycle performance as compared to that of the compressor.

scholarly journals The Aerodynamic Performance of an Over-the-Rotor Liner With Circumferential Grooves on a High Bypass Ratio Turbofan Rotor

2013 ◽  
Author(s):  
Rick Bozak ◽  
Christopher Hughes ◽  
James Buckley
Keyword(s):  
Pressure Ratio ◽  
Axial Compressor ◽  
Aerodynamic Performance ◽  
Oscillating Flow ◽  
Acoustic Liners ◽  
Performance Loss ◽  
Acoustic Liner ◽  
Blade Tip ◽  
Fan Blade ◽  
The Impact

While liners have been utilized throughout turbofan ducts to attenuate fan noise, additional attenuation is obtainable by placing an acoustic liner over-the-rotor. Previous experiments have shown significant fan performance losses when acoustic liners are installed over-the-rotor. The fan blades induce an oscillating flow in the acoustic liners which results in a performance loss near the blade tip. An over-the-rotor liner was designed with circumferential grooves between the fan blade tips and the acoustic liner to reduce the oscillating flow in the acoustic liner. An experiment was conducted in the W-8 Single-Stage Axial Compressor Facility at NASA Glenn Research Center on a 1.5 pressure ratio fan to evaluate the impact of this over-the-rotor treatment design on fan aerodynamic performance. The addition of a circumferentially grooved over-the-rotor design between the fan blades and the acoustic liner reduced the performance loss, in terms of fan adiabatic efficiency, to less than 1% which is within the repeatability of this experiment.

Effect of Blade Passage Surface Heat Extraction on Axial Compressor Performance

2005 ◽  
Author(s):  
P. N. Shah ◽  
C. S. Tan
Keyword(s):  
Heat Transfer ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Heat Extraction ◽  
Transfer Rates ◽  
Blade Passage ◽  
Thermodynamic Effect ◽  
Compressor Performance ◽  
Multistage Compressor ◽  
Corrected Flow

Axial compressor performance with heat extraction via blade passage surfaces (compressor cooling) is compared to its adiabatic counterpart, using computational experiments and meanline modeling. For a multistage compressor with an adiabatic design point, results at fixed corrected rotor speed indicate that if available, compressor cooling would: (i) raise the overall pressure ratio (at a given corrected flow), (ii) raise the maximum mass flow capability, (iii) raise the efficiency, defined as the ratio of isentropic work for a given pressure ratio to actual shaft work, and (iv) provide rear stage choking relief at low corrected speed. In addition, it is found that, if available, cooling in the front stages is better than in the rear stages. This is primarily a thermodynamic effect that results from the fact that, for a given gas, the compression work required to achieve a given pressure ratio decreases as the gas becomes colder. Heat transfer considerations indicate that the engineering challenges lie in achieving high enough heat transfer rates to provide significant impact to the compressor’s performance.

Installation Characteristics of Variable Cycle Engine Based on Inlet Flow Matching

2019 ◽  
Vol 0 (0) ◽  
Author(s):  
Haoying Chen ◽  
Haibo Zhang ◽  
Yong Wang ◽  
Qiangang Zhen
Keyword(s):  
Mathematical Model ◽  
Fuel Consumption ◽  
Geometry Optimization ◽  
Engine Performance ◽  
System Model ◽  
Variable Geometry ◽  
Inlet System ◽  
Supersonic Inlet ◽  
Analysis Scheme ◽  
The Impact

Abstract As per few investigation in installed performance for variable cycle engines, an analysis scheme is proposed on the basis of integrating variable cycle engine and supersonic inlet system model. An integrated mathematical model, containing the inlet and the variable cycle engine is built, realizing the simulation of influences on the installation performance by varying geometry components. The impact on engine performance of variable geometric regulation was analyzed and concluded respectively. The experimental results show that the overflow resistance of the variable cycle engine with variable geometry optimization is reduced at subsonic cruise stage, and the installed fuel consumption is reduced, which significantly improves the installation performance.

Three-Dimensional Aerodynamic Optimization of a Multi-Stage Axial Compressor

2016 ◽  
Author(s):  
Tao Ning ◽  
Chun-wei Gu ◽  
Xiao-tang Li ◽  
Tai-qiu Liu
Keyword(s):  
Genetic Algorithm ◽  
Surrogate Model ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Optimization Method ◽  
Operating Point ◽  
Design Parameters ◽  
Aerodynamic Optimization ◽  
Stall Margin

An optimization method combined of a genetic algorithm, an artificial neural network, a CFD solver and a blade generator, is developed in this research and applied in the three-dimensional blading design of a newly designed highly-loaded 5-stage axial compressor. The adaptive probabilities of crossover and mutation, non-uniform mutation operator and elitism operator are employed to improve the convergence of the genetic algorithm. Considering both the optimization efficiency and effectiveness, a mixture of high-fidelity multistage CFD method and approximate surrogate model of the feed-forward ANN is used to evaluate the fitness. In particular, the database is updated dynamically and used to re-train the surrogate model of ANN for improving the accuracy for predicting. The last stator of the compressor is optimized at the near stall operating point. The tip bow with relative bow height Hb and bow angle αb are treated as design parameters. The adiabatic efficiency as well as the penalty of mass flow and total pressure ratio constitute the objective functions to be maximized. The optimum (Hb = 0.881, αb = 14.7°) obtains 0.4% adiabatic efficiency increase for the whole compressor at the optimized operating point. The detailed aerodynamic is compared between the baseline and optimized stator, and the mechanism is analyzed. The optimized version obtains 5.1% increase in stall margin and maintains the efficiency at the design point.

Impact of Variable-Geometry Stator Hub Leakage in a Low Speed Axial Compressor

1998 ◽  
Author(s):  
Wendy S. Barankiewicz ◽  
Michael D. Hathaway
Keyword(s):  
Mechanical Design ◽  
Axial Compressor ◽  
Leading Edge ◽  
Operating Conditions ◽  
Variable Geometry ◽  
Turning Angle ◽  
Blade Loading ◽  
Axial Compressors ◽  
The Impact ◽  
Variable Stator

The impact of hub leakage flow associated with the clearance gaps of hub-shrouded variable-geometry stator rows in axial compressors is investigated experimentally. The objectives of this work are to investigate the sensitivity of performance to chordwise leakage location and to provide guidance for the mechanical design of variable stator hub trunions. Although blade loading near the hub is increased when leakage occurs at the leading edge, losses also increase for both design and off-design operating conditions. Leading edge leakage also causes a greater spanwise variation in absolute turning angle over the first 20% of span from the hub. Results show that for a moderately separated stator, the optimum leakage configuration features trailing edge leakage with the leading edge sealed. This confirms the current practice of most engine companies in placing the hub trunion at the blade leading edge.

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