Impact of Uniform Surface Roughness on the Aerodynamic Performance of an Axial Compressor Rotor at Different Rotational Speeds

The aerodynamic performance of a compressor rotor is known to deteriorate due to surface roughness. It is important to understand this deterioration as it impacts the overall performance of the engine. This paper, therefore, aims to numerically investigate the impact of roughness on the performance of an axial compressor rotor at different rotational speeds. In this numerical study, the simulations are carried out for NASA Rotor37 at 100%, 80%, and 60% of its design speed. with and without roughness on the blade surface. These speeds are chosen because they represent different flow regimes. The front stages of a multistage compressor usually have a supersonic or transonic regime whereas the middle and aft stages have a subsonic regime. Thus, these performance characteristics can give an estimate of the impact on the performance of a multistage compressor. At 100% speed (design speed), the relative flow is supersonic, at 80% of design speed, the relative flow is transonic and at 60% of design speed, the relative flow is subsonic. Detailed flow field investigations are carried out to understand the underlying flow physics. The results indicate that, for the same amount of roughness, the degradation in the performance is maximum at 100% speed where the rotor is supersonic, while the impact is minimum at 60% speed where the rotor is subsonic. Thus, the rotor shock system plays an important role in determining the performance loss due to roughness. It is also observed that the loss increases with increased span for 100% and 80% speeds, but for 60% speed, the loss is almost constant from the hub to the shroud. This is because, with the increased span, the shock strength increases for 100% and 80% speeds, whereas at 60% speed flow is subsonic.

Periodic Unsteady Kinematics of Hub Flows in a Shrouded Multistage Compressor

Periodic unsteady flow kinematics in a shrouded multistage low-speed axial compressor has been measured for the first time. Data have been acquired at the inlet and exit of a shrouded 3rd- stage stator with a particular focus on the hub flows. The newly found features of the hub flow in a shrouded multistage compressor are different from those at the midspan or in unshrouded (i.e., cantilevered) compressors. First, the merging of the 2nd-stage stator and 3rd-stage rotor wakes causes positive radial migration near the rotor wake pressure surface at the hub of the 3rd-stage stator inlet. Second, the low-momentum labyrinth seal leakage flow of the 3rd-stage stator merges with the 3rd-stage rotor wake to create streamwise vorticity at the 3rd-stage stator inlet hub. Third, contrary to unshrouded stators, suction side hub corner separation in the shrouded 3rd-stage stator reduces rotor wake stretching. Thus, velocity disturbances are attenuated less, and amplitudes of periodic fluctuations in flow angles are larger at the 3rd-stage stator exit hub than at midspan. The positive radial migration of the rotor wake hub flow and wake stretching reduction are expected to decrease efficiency, whereas streamwise vorticity generation is expected to increase efficiency.

Impact Investigation of Stator Seal Leakage on Aerodynamic Performance of Multistage Compressor

In this paper, a numerical model based on the mass flow rate of seal leakage is presented, and a 3D numerical method of a multistage axial compressor with good engineering practicability is established. Validation consists of modeling a nine-stage axial compressor in all operating rotation speeds and calculating results of the performance characteristic curves in good agreement with test data. Comparisons are made against different cases of seal leakage mass flow rate for analyzing the impact of increasing seal leakage on the aerodynamic performance of the multistage axial compressor. The results indicate that the performance of the nine-stage axial compressor is degenerated faster and faster with seal leakage increasing in all operating working points, and the degeneration of performance of this compressor can be evaluated by the relationships of main performance parameters with the mass flow rate of seal leakage. Comparisons of flow distribution in the compressor for different cases of seal leakage also show that stators located in front stages of the multistage axial compressor are affected more seriously by the increasing seal leakage, and it can be confirmed that relatively larger flow losses in front stages bring significant impact on the decay of aerodynamic performance of a multistage axial compressor.

Compressor Maps and Coupling: Symmetry, Paradox, and Clarity

The most common compressor map framework, referred to here as the β-framework, will be shown to suffer from limitations that grow more troublesome in the multiple-map environment. When maps are coupled in series in the β-framework, it is very common to find operating points that are physically unrealizable, but these cannot generally be avoided without first generating them. A feasible situation is described in which the β-framework leads to an apparent physical paradox. In the proposed S-framework, the map itself is recast in terms of independent variables (corrected speed and exit corrected flow) and dependent variables (inlet corrected flow and temperature ratio). The propagation of information in map coupling is split into an upstream-marching corrected flow ‘flux’ and a downstream-marching temperature ‘flux’. Finding the equilibrium operating point requires only finding a simple intersection between curves. The S-framework is then developed further into a more compact S’-framework that exhibits a natural set of qualitative symmetries. The S- and S’-frameworks are shown to simplify compressor map expression, resolve the problems shown with the β-framework, and aid intuition with regard to off-design phenomena. The resolution of the paradox using the S’-framework is a new description of multistage compressor performance hysteresis.

Periodic Unsteady Kinematics of Hub Flows in a Shrouded Multistage Compressor

Periodic unsteady flow kinematics in a shrouded multistage low-speed axial compressor has been measured for the first time. Data have been acquired at the inlet and exit of a shrouded 3rd-stage stator with a particular focus on the hub flows. The newly found features of the hub flow in a shrouded multistage compressor are different from those at the midspan or in unshrouded (i.e., cantilevered) compressors. First, the merging of the 2nd-stage stator and 3rd-stage rotor wakes causes positive radial migration near the rotor wake pressure surface at the hub of the 3rd-stage stator inlet. Second, the low-momentum labyrinth seal leakage flow of the 3rd-stage stator merges with the 3rd-stage rotor wake to create streamwise vorticity at the 3rd-stage stator inlet hub. Third, contrary to unshrouded stators, suction side hub corner separation in the shrouded 3rd-stage stator reduces rotor wake stretching. Thus, velocity disturbances are attenuated less, and amplitudes of periodic fluctuations in flow angles are larger at the 3rd-stage stator exit hub than at midspan. The positive radial migration of the rotor wake hub flow and wake stretching reduction are expected to decrease efficiency, whereas streamwise vorticity generation is expected to increase efficiency.

Investigation of Seal Cavity Leakage Flow Effect on Multistage Axial Compressor Aerodynamic Performance with a Circumferentially Averaged Method

The seal cavity leakage flow has a considerable impact on the performance of the aeroengine, especially on the multistage compressor. Thus, a quasi-three-dimensional simulation program named CAM is developed basing on circumferentially averaged throughflow method. The program enables a rapid diagnosis for the performance degradation of multistage compressor caused by labyrinth wear. The coupling flow field between the seal cavity leakage flow and the main flow field at the root of the shrouded stator of a high-loading three-stage compressor with inlet guide vanes (IGV) was simulated by CAM and the results indicate that seal cavity leakage flow has a significant impact on the overall performance of the compressor. That is, for a 1% increase in the seal-tooth clearance-to-span ratio, the decrease in total pressure ratio was 2.6%, and the reduction in efficiency was 0.6%. Stage performance shows that the seal cavity leakage flow reduces the pressurization capacity of the current stator and the work capacity of the downstream rotor, but has little effect on the upstream blade row. Spanwise distribution of blade element performance shows that the leakage flow leads to an increased flow blockage near the hub, resulting in spanwise migration. The incidence of the stator and rear rotor then change through the entire span. The leakage flow leads to the flow blockage and migration and hence changes the incidence angle, which results in the deterioration of compressor performance.

Dust Ingestion in a Rotorcraft Engine Compressor: Experimental and Numerical Study of the Fouling Rate

Helicopters often operate in dusty sites, ingesting huge amounts of contaminants during landing, take-off, hover-taxi, and ground operations. In specific locations, the downwash of the rotor may spread soil particles from the ground into the environment and, once ingested by the engine, may stick to the compressor airfoils. In the present work, the Allison 250 C18 engine’s multistage axial-flow compressor is employed to study the fouling rate on rotor blades and stator vanes from both numerical and experimental standpoints. The compressor is operated in a typical ground-idle operation, in terms of the rotational regime and contaminant concentration, in laboratory-controlled conditions. The mass of deposits is collected from the airfoil surfaces at the end of the test and compared to that estimated through the numerical model. The experimental test shows that the airfoils collect almost 1.6% of the engine’s total mass ingested during a ground-idle operation. The capability of numerical methods to predict the fouling rate on the rotating and stationary airfoils of a multistage compressor is tested through the implementation of literature based deposition models. Sticking models show a good agreement in terms of the relative results; nevertheless, an overestimation of the deposited mass predicted is observed.

Interstage Pressures of a Multistage Compressor with Intercooling

This paper considers the criterion of minimum compression work to derive an expression for the interstage pressure of a multistage compressor with intercooling that includes the gas properties, pressure drops in the intercoolers, different suction gas temperatures, and isentropic efficiencies in each compression stage. The analytical expression for the interstage pressures is applied to estimate the number of compression stages and to evaluate its applicability in order to estimate interstage pressures in the operation of multistage compressors, which can be especially useful when their measurements are not available.

Deposition Pattern Analysis On a Fouled Multistage Test Compressor

Compressor fouling is one of the main causes of gas turbine performance degradation. Microsized particles adhere to the blade surfaces increasing the surface roughness and modifying the airfoil shape. In the present work, the contamination of the Allison 250 C18 multistage compressor engine with four sorts of micrometric dust has been provided. The tests were performed changing the relative humidity at the compressor inlet and the unit rotational speed. After each test, a photographic inspection of the internal fouled parts has been realized and the digital pictures have been analyzed employing an image processing package. The deposits build-up of stator vanes and rotor blades have been postprocessed and the most affected regions of each compressor stage have been highlighted. Besides, a numerical simulation of the machine has been performed. The numerical flow field has been used to highlight the blade regions which show the most favorable conditions for particle deposition. A theoretical model has been applied to the flow field to simulate the particle deposition. The combination of the deposition model with the results of the CFD simulations gives the chance to better understand the experimentally-founded deposition patterns. Those results have been finally compared to the pictures of the patterns. The possibility to detect and measure the deposition patterns on a rotating test rig and the comparison with models and experiments gave the possibility to assess in detail the particle deposition phenomenon on a multistage axial compressor flow path.

The Investigation of Axial Compressor With Aggressive Flow Path Near Stall Condition

Rotating stall is an unsteady flow phenomenon in flow rate reduction process of axial compressor. Single or multiple stall cells can be found in blade passage, rotating circumferentially at frequency ranging from 20% to 80% of compressor rotating frequency. Rotating stall may further lead to surge, then mass flow and pressure ratio of the compressor decrease rapidly while vibration amplitude of bearing fulcrum increases significantly, which would affect safety operation of the aero-engine. This paper reports the study on a multistage compressor with special-shaped aggressive flow path, which is significantly different from the conventional compressor with its meridian flow path midline parallel or nearly parallel to the rotating shaft axis. To study rotating stall and surge characteristics of multistage compressor with inlet distortion, this paper takes a 4.5-stage axial compressor as research object which has special S-shaped meridian flow path. On SJTU compressor test facility, the author carries out multidisciplinary tests on compressor dynamic performance (stator LE/TE and inter-stage fluctuating pressure measured by fluctuating pressure transducers, vibration measured by acceleration transducers, and dynamic stress measured by pasting strain transducers on the surface of stator blade). And in subsequent surge experiment, stall and surge signals at different rotating speeds are successfully captured. With above test data, axial, radial and circumferential instability development characteristics are studied by time domain correlation analysis, frequency domain power spectrum and coherence analysis, which preliminarily reveals the generation, development, propagation and elimination characteristics of rotating stall and surge. Meanwhile, in view of time-frequency characteristics of two instability transition processes, time-frequency analysis method based on Choi-Williams distribution model is adopted to capture instantaneous frequency characteristics. The results show that instability phenomenon exhibits quasi-exponential change in instability transition processes. Compared with traditional methods, time-frequency method is more effective in characterizing strong time-varying instability signals such as rotating stall and surge.