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The rotor blade tip leakage flow and associated formation of the tip leakage vortex and interaction of the tip leakage vortex with the shockwave, particularly in the case of a transonic compressor rotor have significant impact on the compressor performance and its stability. Air injection upstream of the compressor rotor tip has been shown to improve compressor performance and enhance its stability. The air required for rotor blade tip injection is generally taken from the later stages of the compressor thus causing penalty on the gas turbine performance. In this study, effects of water injection at the rotor tip with and without the wet compression on the compressor performance and its stability have been examined. To achieve the stated objectives, the well tested transonic compressor rotor stage, NASA rotor stage 37, has been numerically simulated. The evaluation of results on various performance parameters, such as total pressure ratio, inlet flow capacity, and adiabatic efficiency combined with contours of total pressure losses, entropy, Mach number, and temperature including limiting streamlines, shows that the blade tip water injection could help in reducing low energy region downstream of the shockwave and strength of the tip leakage vortex with the compressor operating at its rotating stall boundary condition. The extent of reduction depends on the droplet size, injection flow rate, and its velocity. Furthermore, results show that combined case of the blade tip water injection and the wet compression could provide better stall margin enhancement than the blade tip water injection case.

The current paper first reviews experimental and numerical investigations to understand flow physics and to develop optimum configurations of circumferential grooves in compressor rotors. Circumferential grooves are used mainly to increase stall margin in axial compressors with small decrease in aerodynamic efficiency. Although circumferential groove casing treatment has been used widely, flow mechanisms of the circumferential grooves at near stall conditions are not well understood yet. Detailed time-dependent flow measurement inside tip gap in a high speed compressor is still a big challenge even though significant advance has been made in non-intrusive flow measurement technique. Therefore numerical approaches have been used to study relevant flow physics. However, optimum design of circumferential grooves to a given compressor with the computational tools is not practical yet. In the present paper, various investigations to study flow physics of circumferential groove casing treatment in axial compressor are reviewed first. Possible missing flow physics are identified and future research efforts for the optimum design are discussed.

Transonic compressor rotor performance is sensitive to variations in several known design parameters. One such parameter is the chordwise location of maximum thickness. This article reports on the design and experimental evaluation of two versions of a low aspect ratio transonic rotor that had the location of the tip blade section maximum thickness moved forward in two increments from the nominal 70 percent to 55 and 40 percent chord length, respectively. The original hub characteristics were preserved and the maximum thickness location was adjusted proportionately along the span. Although designed to satisfy identical design speed requirements, the experimental results reveal significant variation in the performance of the rotors. At design speed, the rotor with its maximum thickness located at 55 percent chord length attains the highest peak efficiency among the three rotors but has lowest flow rollback relative to the other two versions. To focus on current ruggedization issues for transonic blading (e.g., bird and ice ingestion), detailed comparison of test data and analysis to characterize the aerodynamic flow details responsible for the measured performance differences were confined to the two rotors with the most forward location of maximum thickness. A three-dimensional viscous flow analysis was used to identify the performance-enhancing features of the higher efficiency rotor and to provide guidance in the interpretation of the experimental measurements. The computational results of the viscous analysis show that the difference in performance between the two rotors can be attributed to the higher shock losses that result from the increased leading edge “wedge angle” as the maximum thickness is moved closer to the leading edge. The test data and the three-dimensional viscous analysis also reveal that the higher efficiency rotor achieves the same static pressure rise potential and loading at a higher flow level than its less efficient counterpart and this is responsible for its resulting lower flow rollback and apparent loss in stall margin. Comparison of the peak efficiencies attained by the two rotors described in this article with the baseline ruggedized rotor performance presented in part 1 of this paper suggests the existence of an optimum maximum thickness location at 55 to 60 percent chord length for such low aspect ratio transonic rotors.

The importance of the unsteady effects between the casing treatment and rotor is still not clear. Experiments are conducted in a transonic compressor with arc skewed slot casing treatment configurations. The experimental results indicate that the unsteady excitation frequency due to the slot number of casing treatment is one of the most important factors influencing compressor performance. Also, compressor performance can be overall enhanced through optimizing this unsteady excitation frequency. For the transonic compressor herein, peak efficiency, stall margin, and maximum flow mass can be improved by 0.17%, 19.86%, and 0.81%, respectively, at near design rotating speed, which can reach up to 1.13%, 57.84%, and 1.57%, respectively, at part design speed.

Sand erosion is a phenomenon whereby solid particles impinging on a wall cause serious mechanical damage to the wall surface. Aircraft engines operating in a particulate environment are subjected to deterioration of performance and life due to sand erosion. In particular, the compressors of aircraft engines are severely affected. In our previous study, the authors developed code for predicting sand erosion three-dimensionally and subsequently applied it to a subsonic axial compressor. However, design trends for modern compressors continue to increase blade loading and decrease aspect ratios. Additionally, modern compressors are transonic. To aid in the design and deepen understanding, it is important to study sand erosion in such transonic compressors. Herein, Rotor 37 was used in sand erosion simulations. The particle motion, erosion rate were investigated.Copyright © 2010 by ASME

This contribution is aimed at summarizing the effort taken to apply stereoscopic PIV (SPIV) measurements in the tip clearance of a transonic compressor rotor equipped with a casing treatment. A light sheet probe was placed downstream of the stator and aligned to pass the light sheet through a stator passage into the blade tip clearance of the rotor. A setup with three cameras has been used in order to record the entire 2C velocity field and the smaller area of 3C field of view at the same time instance for comparison with earlier 2C PIV results. A homogeneous seeding distribution was achieved by means of a smoke generator, working with evaporated and subsequently re-condensed mineral oil. The main emphasis of the SPIV measurement was to establish a data set with high spatial resolution close to the compressor casing, where the aerodynamic effects of a CT are known to be strong. Additionally the SPIV data was intended to be used for comparison with numerical results and related code validation, as detailed laser-based flow field investigations in the vicinity of a casing treatment in a transonic compressor environment are barely reported in the literature. The paper will discuss some major aspects of the utilized PIV data processing and point out a variety of frequently underestimated error sources that influence the overall quality of the recovered data in spite of the fact that the individual PIV recordings seemed to be of very good quality. Thus the authors will not focus on the PIV results and related interpretation of the flow field, but on the optimization and procedures applied during setup of the experiment and data processing respectively.

An investigation of the behavior of a transonic compressor rotor when operating close to stall is presented. The specific areas of interest are the behavior and location of lowfrequency instabilities close to stall. In running close to stall, compressors can begin to exhibit nonperiodic flow between the blade passages even when appearing to be operating in a stable steady-state condition. The data from the current rotor clearly show that low-frequency instabilities were present during steady-state operation when stall was approached. These frequencies are not geometrically fixed to the rotor and typically appear at 0.3–0.8 of the rotor speed. The presence of these low-frequency instabilities is known and detection is reasonably commonplace; however, attempts to quantify the location and strength of these instabilities as stall is approached have proved difficult. In the current test fast response pressure sensors were positioned in the case-wall; upstream, downstream, and over the rotor blade tips. Simultaneous data from the sensors were taken at successive steady-state settings with each being closer to stall. A time domain analysis of the data investigates the magnitude of the instabilities and their transient effect on the relative inlet flow angle. The data are also presented in the frequency domain to show the development and distribution of the instabilities over the rotor as stall was approached. Initially the instabilities appeared within the rotor row and extended downstream but at operation closer to stall they began to protrude upstream as well. The greatest amplitude of the instabilities was within the blade row in the complex flow region that contains phenomena such as the tip-vortex/normal-shock interaction and the shock/boundary-layer interaction. In addition as stall is approached the growth of the instabilities is nonlinear and not confined to one frequency. The present study was part of the compressor research program sponsored by the Propulsion and Power Department of the Naval Air Warfare Centre, Patuxent River, MD with Ravi Ravindranath as the technical monitor.

Time-resolved radial transport has been measured in a transonic compressor rotor by injecting a thin sheet of tracer gas upstream of the rotor and then surveying the tracer concentration at the rotor exit. The simultaneous, co-located, high-frequency response measurements of local tracer gas concentration, total temperature, and total pressure made downstream of the rotor showed that most of the fluid transported radially appears in the blade wakes and that this fluid has considerably higher entropy than the circumferential mean. Both inward and outward fluid transport along the span was observed (3.5 percent of the total throughflow moved toward the tip while 1.6 percent moved toward the hub). Tracer concentration and fluid total temperature and pressure varied considerably from wake to wake, even on multiple samplings of the same blade. The time mean spreading rate inferred from these measurements is in general agreement with previously reported studies on multistage low-speed compressors and is well predicted by the method of Gallimore and Cumpsty. It is suggested that a vortex street in the blade wakes could be responsible for both the observed radial transport and the large wake-to- wake variability. A quasi-three-dimensional model of a vortex street wake was developed and shown to be consistent with the data. The model predicts all of the inward transport but only 20 percent of the outward transport. It is hypothesized that outflow in separated regions on the blade suction surface is responsible for the remainder of the transport toward the rotor tip. Since the entropy, as well as the mass of the fluid transported radially, was measured, an estimate of the redistribution of loss in rotor due to radial fluid transport could be made. This showed that the effect of radial transport in this rotor was to move substantial loss from the rotor hub to tip, implying that a conventionally measured spanwise efficiency survey may not accurately represent the performance of individual blade sections.

An investigation of the behavior of a transonic compressor rotor when operating close to stall is presented. The specific areas of interest are the behavior and location of low-frequency instabilities close to stall. In running close to stall, compressors can begin to exhibit nonperiodic flow between the blade passages even when appearing to be operating in a stable steady-state condition. The data from the current rotor clearly show that low-frequency instabilities were present during steady-state operation when stall was approached. These frequencies are not geometrically fixed to the rotor and typically appear at 0.3–0.8 of the rotor speed. The presence of these low-frequency instabilities is known and detection is reasonably commonplace; however, attempts to quantify the location and strength of these instabilities as stall is approached have proved difficult. In the current test fast response pressure sensors were positioned in the case-wall; upstream, downstream, and over the rotor blade tips. Simultaneous data from the sensors were taken at successive steady-state settings with each being closer to stall. A time domain analysis of the data investigates the magnitude of the instabilities and their transient effect on the relative inlet flow angle. The data are also presented in the frequency domain to show the development and distribution of the instabilities over the rotor as stall was approached. Initially the instabilities appeared within the rotor row and extended downstream but at operation closer to stall they began to protrude upstream as well. The greatest amplitude of the instabilities was within the blade row in the complex flow region that contains phenomena such as the tip-vortex/normal-shock interaction and the shock/boundary-layer interaction. In addition as stall is approached the growth of the instabilities is nonlinear and not confined to one frequency.

The loss mechanism due to the blade row interaction effects in a compressor is an important and interesting problem. In the present paper, the loss in a compressor stage under both peak efficiency (PE) and near stall (NS) operating conditions is studied numerically by solving the 3D unsteady RANS equations. The loss is measured in terms of the thermodynamic irreversibility. In order to examine the relationship between loss and flow structures, an intensive parameter called the intensity of irreversibility is derived. The relative importance of various flow structures in loss production has been identified by the orders of the intensity of irreversibility. An audit of the irreversibility has been carried out, which shows the importance of suction surface boundary layers and the tip leakages in loss generation. Under the NS conditions, the tip leakages are the most important sources of loss production. The unsteady evolution of irreversibility in the rotor and stator passage is also analyzed.

Defining the thermodynamic efficiency of the wet compression process in a compressor is not trivial, since the flow in this case has multiple phases present which interact with each other. In this paper, an approach is presented that calculates the overall entropy creation and thus the isentropic efficiency of a wet compression process in a transonic compressor rotor. The viscous dissipation function is calculated everywhere in the domain in the post-processing phase of the CFD simulation and integrated to the wall, with special treatment in the near-wall regions where high rates of entropy generation occur. The isentropic efficiency of the wet compression is then determined from the entropy generation rate. Analytical integration of wall functions and numerical integration of the viscous dissipation function allows for reasonable results even with relatively coarse grids and can be applied for single-phase flows. The methodology presented is also useful to quantify the efficiency of thermodynamic processes in devices that introduce streams into the flow path, such as cooled turbines and compressors with flow control.

This article presents the study of casing treatments on an axial compressor stage for improving stability and enhancing stall margin. So far, many simulations of casing treatments on single rotor or rotor-stator configurations were performed. But as the application of casing treatments in engines will be in a multistage compressor, in this study, the axial slots are applied to a typical transonic first stage of a high-pressure 4.5-stage compressor including an upstream IGV, rotor, and stator. The unsteady simulations are performed with a three-dimensional time accurate Favre-averaged Navier-stokes flow solver. In order to resolve all important flow mechanisms appearing through the use of casing treatments, a computational multiblock grid consisting of approximately 2.4 million nodes was used for the simulations. The configurations include axial slots in 4 different variations with an axial extension ranging into the blade passage of the IGV. Their shape is semicircular with no inclination in circumferential direction. The simulations proved the effectiveness of casing treatments with an upstream stator. However, the results also showed that the slots have to be carefully positioned relative to the stator location.

A numerical investigation of the NASA Rotor 35 transonic compressor rotor at near stall conditions was conducted with an unsteady RANS technique. A series of time-step size and inner iterations were chosen to create a design-of-experiment matrix to study their impact on the unsteady flow field. Solution accuracy is determined by comparisons with LDV, and pressure and temperature profiles. Results obtained showed that for this near stall unsteady flow field unsteady analysis is required. However, an asymptotic relationship does not exist between increased numerical temporal accuracy and improved solution predictions. This paper discusses the physical flow features of a transonic rotor near stall as well as reasons why increased temporal accuracy does not guarantee the solution to asymptotically approach the data.Copyright © 2008 by ASME

In order to accurately model the physics associated with losses in a transonic compressor, a time-accurate simulation of a transonic compressor rig was developed. Initially phase-lag simulations were conducted to determine required time stepping and grid re nement to capture the blade row interactions. These parameters were used to develop quarter annulus simulations to investigate the physics involved in the rotor bow shock interaction with a highly loaded upstream blade row and its e ect on the compressor. Three di erent axial spacings between the rotor and the upstream blade row have been simulated corresponding to the experimental rig. Details of the unsteady and time-averaged ow eld are presented and described. Among the many ow features resolved in the simulations was a suction side separation bubble in the upstream stator blade row that forms and collapses as a result of di usion and passing pressure waves.

To improve the understanding of how sweep affects the 3D shock structure and the shock-vortex interaction, two transonic compressor bladings are measured with a 3D Laser-2-Focus setup, and the flow is analyzed for comparable operating conditions. Rotor 1 features radially stacked profiles, while rotor 3 introduces forward sweep. Performance comparisons of these rotors [1] indicated that forward sweep, as introduced in this design, has a beneficial effect on performance and stall margin. The detailed experimental analysis of the 3D shock structure and the influence of tip leakage vortex interaction at design speed is the scope of this paper.

In the present paper, the flow structure inside a low-solidity diffuser of a transonic compressor was investigated in detail. Steady computations were carried out and compared to experimental data. The secondary flow development inside the diffuser was analyzed and the reason for the stall inception was detected. Unsteady calculations were performed for two operating points, one close to the choke and the other one close to the stall of the compressor, in order to assess the effect of the unsteadiness in the diffuser secondary flow development.Copyright © 2002 by ASME

The information provided in this report is a summary of the research activities completed during the period of July 1974 to January 2002. Complete details on these research activities are provided in the 48 technical reports and 26 journal articles that were generated throughout the effort. A listing of these reports and journal articles are provided in the appendix of this summary. Seven distinct research efforts (all documented by the reports and journal articles identified in the appendix) were pursued during the reporting period.

This paper investigates the flutter mechanism of a high loaded transonic compressor rotor by applying different numerical methods. For predictions of the stability in the frequency domain, harmonic fluid forces are calculated with a linearized Euler method and with a nonlinear time domain Navier-Stokes solver for forced motion. To include nonlinear effects in the stability prediction, the time domain method is applied in a fluid-structure coupled mode of operation. The results of this 3D viscous aeroelastic computations in the time domain for the whole annulus of the rotor are compared to the frequency domain solutions.

A set of inlet guide vane (IGV) unsteady surface pressure measurements is presented. The unsteady aerodynamic effects of a highly loaded, high speed downstream compression stage on the upstream inlet guide vane stator surface pressures are characterized through experimental and computational analysis. The axial spacing between the IGV and rotor was varied between 12%, 26%, and 56% of the IGV chord for a 105% speed, peak efficiency operating condition, which is transonic. Unsteady IGV surface pressures were acquired for two spanwise locations on both blade surfaces. The largest unsteady surface pressure magnitudes were obtained at the 12% axial spacing configuration and 95% chord location. In general, spanwise variations were found to be important. The upstream bow shock effect is non-linear in character. Comparisons to a two-dimensional, non-linear unsteady multi-blade row Navier-Stokes analysis at 50% span show a good agreement with the IGV unsteady surface pressure results and higher harmonic content. The results of the study indicate significant variations in the IGV unsteady loading caused by changes in axial spacing.