Performance Of Multistage Axial Compressor Research Articles

Aerodynamic Performance of Multi-stage Axial Compressor Considering Cold-to-Hot Deformation Effect

Predicting the exact performance of multi-stage compressors for an aircraft gas turbine is important in the design process and calculation of off-design characteristics. This study explores the quantitative effects of cold-to-hot deformation on the aerodynamic performance of a three-stage axial compressor in terms of power consumption and efficiency. The reason of difference between numerical result and performance test data is investigated. Three-dimensional computational fluid dynamics (3D CFD) was used to predict the flow characteristics. A finite-element (FE) analysis was conducted to calculate the displacement caused by rotation, thermal expansion, and pressure force. First, the fluid domain of the compressor with cold geometry was analyzed via a 3D CFD analysis, and the displacement for all coordinates was then obtained from the FE analysis. The hot geometry and mesh for the fluid domain were generated based on the cold geometry and calculated displacement. Finally, the aerodynamic performance of the compressor with hot geometry was evaluated via a 3D CFD analysis. The results for the hot geometry compressor were more consistent with the performance test data than those for the cold geometry compressor in terms of the pressure ratio and efficiency based on the mass flow rate. Moreover, the average rotor tip clearance decreased by approximately 0.5% with respect to the span height by the cold-to-hot deformation. The total pressure loss at the rotor tip region was lower for the hot geometry compressor than the cold geometry compressor because of the reduced entropy of the high vortex loss core around the rotor tips. The results showed that the power consumption was 1.16% higher for the hot geometry compressor than the cold geometry compressor at the designed pressure ratio, which was primarily caused by an increase in the mass flow rate due to the wider throat area after the cold-to-hot deformation.

An 11-stage axial compressor performance simulation considering the change of tip clearance in different operating conditions

The multistage axial compressor is one of the critical components of aero-engines and plays a key role in their performance, reliability, and economy. Tip clearance has a significant impact on the performance and stability of multistage axial compressors. Due to blade and disk deformations, tip clearance will vary significantly in different operating conditions. Thus, tip clearance should be accurately estimated when evaluating compressor performance. This paper proposes a new model to predict changes of tip clearance of multistage axial compressors in different operating conditions. A first-principles approach is used to estimate the change of tip clearance caused by thermal and mechanical deformation. The span-wise temperature distribution across each stage of the multistage compressors is considered by the proposed model in this paper. The model was validated by General Electric Company (GE) E3 engine experimental results. Using the model, the performance of an 11-stage axial compressor is simulated. The results show that accounting for tip clearance variations has a 0.5% impact on the calculated mass flow rate and a 1% impact on the calculated efficiency. Thus, variations of tip clearance at different operating conditions cannot be ignored and the proposed new model is useful to accurately predict the performance of multistage axial compressor.

Axisymmetrically Stalled Flow Performance for Multistage Axial Compressors

A study of the stalled flow performance of multistage axial compressors is presented. A proposal is made regarding the form of axisymmetric pumping performance in stall (which is a requisite of current rotating stall models) over the entire compressor flow range, including reversed flow. It is also shown that the axisymmetric performance can rise above the measured stall point pressure rise, thus indicating greater unstalled pressure rise potential. A simple two-dimensional reversed flow model is presented, and is shown to be in reasonable agreement with available high backflow compressor data. The model predicts that the blade stagger angle greatly influences the reversed flow characteristic. Calculations are also carried out applying this axisymmetric characteristic to the rotating stall model of Moore.

An Axial Compressor End-Wall Boundary Layer Theory

The essential ingredient missing in existing prediction methods for the performance of multistage axial compressors is that which would account for the effect of end-wall boundary layers. It is, in fact, believed that end-wall boundary layers play a major role in compressor performance and the absence of an adequate theory represents a handicap to turbomachinery designers that might be likened to the handicap that designers of wings, for example, would face if Prandtl had not introduced the idea of a boundary layer. In this paper a new theory is developed which retains all elements of classical boundary layer theory; for example, we discuss variables such as momentum thickness and wall shear stress. However, the present theory introduces new concepts such as axial and tangential defect force thickness, a rotor exit-stator inlet “jump condition” and the importance of these concepts is demonstrated. Inherent in the derivation is an identification of the role of secondary flow and tip clearance flow. A proper means of matching the boundary layer calculations to conventional main stream calculations is suggested. Independent of empirical parametization it appears that the theory is capable of correctly modeling boundary layer blockage, losses, and end-wall stall. Near stall, the main stream-boundary layer interaction is very strong.

An Improved Axial Flow Compressor Design

In using the method of stage stacking to compute the off-design performance of multi-stage axial compressors, it has been observed that the limitation on performance at speeds above the design speed has been set by the stall and the choke points of the rear stages(1). Thus if the rear stages can absorb a wide range of mass flows between stalled conditions and choked conditions, a better performance could be obtained.Compressor stages using low stagger blades will absorb a large range of mass flow between stalled and choked condition; but because of the high axial velocity involved in their use, they tend to be unsuitable for low pressure stages because of the high Mach number obtained. In the higher pressure stages the increased gas temperature will lower the Mach number for the same velocity and give more efficient operation.

Use of Experimental Interstage Performance Data to Obtain Optimum Performance of Multistage Axial Compressors

Experimental interstage performance data provide the most useful basis currently available for correction of faults in performance of multistage compressors. However, to describe stage performance adequately requires much more information than the amount needed to describe over-all compressor performance, and to interpret the data requires that all of this information be correlated. Because the wealth of the data may conceal its meaning, some simple method is needed to focus attention on the matching of one stage to another and to the over-all compressor as well. A technique is presented whereby performance of individual stages can be depicted graphically upon the performance map of an over-all multistage compressor. Through this presentation the individual and over-all effects can be examined together. Insight is afforded into performance relationships of one stage to another; and faults can be diagnosed if any exist in the matching, radial or axial, of individual stages. Changes of blade geometry can then be prescribed upon a clear and explicit basis. To illustrate use of the technique, data are presented describing gains that have been achieved experimentally.