ОПРЕДЕЛЕНИЕ ДРОССЕЛЬНОЙ ХАРАКТЕРИСТИКИ ТУРБОВАЛЬНОГО ГТД НА ОСНОВЕ МЕТОДА МАТЕМАТИЧЕСКОГО МОДЕЛИРОВАНИЯ С ИСПОЛЬЗОВАНИЕМ ОДНО- И ДВУМЕРНЫХ ПОДХОДОВ К РАСЧЕТУ ПАРАМЕТРОВ КОМПРЕССОРА

The internal losses in the tip clearance region strongly influence the compressor performance and its operational range. Previous research proved that passive wall treatments with circumferential grooves in axial compressors effectively increase the compressor stall margin. The vortex generated inside the circumferential grooves creates a resistance to the flow that leaks into the tip clearance region of the compressor. However, most works found in the literature on circumferential grooves in axial compressors deal only with high-performance single-stage axial compressors. Therefore, there is a need to investigate and analyze the behavior of circumferential grooves in a multi-stage environment. In the present work, a passive wall treatment with circumferential grooves was implemented in a multi-stage axial compressor. Different configurations of circumferential grooves were created at the casing of the first and second rotor rows used in a four-stage axial flow compressor. Numerical simulations were performed to evaluate the influence of the circumferential grooves on the performance of a multi-stage axial compressor. The results obtained after the simulations for the different circumferential groove configurations were compared with the results obtained for the compressor without casing treatment (smooth wall) for different rotational speeds. Furthermore, the complete compressor map characteristics were simulated for the different casing treatment configurations, and the results were compared with the compressor characteristics of the smooth wall case. The passive wall treatment with circumferential grooves produced changes in the multi-stage axial compressor flow field, especially in the tip clearance region, improving the compressor stability mainly for part load speeds.

A Multi-Stage Axial Compressor Characteristic Calibration Method based on One-Dimensional Average Streamline Method

The operational envelope of gas turbine engines such as employed in the Army Blackhawk helicopter is constrained by the stability limit of the compression system. Stall control technologies developed at the ARL Vehicle Technology Directorate and NASA Glenn Research Center to improve the stable operating range of compressors lack fundamental understanding beneficial to design guidance. Improved understanding of the stall inception process and how stall control technologies mitigate such will provide compressors with increased tolerance to stall, thereby expanding the operational envelope of military gas turbine engines. To provide improved understanding, time accurate CFD simulations are being computed at full-scale actual engine conditions using the parallel solver, TURBO, developed at Mississippi State University (MSU) with support from NASA, DoD, and industry. As a compressor nears stall the flow field is no longer periodic from passage to passage, therefore full-annulus (i.e., all blade passages of every blade row) must be computed. For a typical multistage compressor this becomes a formidable computational challenge requiring access to massively parallel machines in order to meet the computational and memory demands of the problem. Unsteady full-annulus simulations of both a single stage axial and centrifugal compressor, for which detailed experimental data are available, provide validation of the TURBO code for subsequent use in predicting the stall inception processes in an axicentrifugal multistage compressor employed in the Army Blackhawk helicopter. The simulations show the inception of stall and subsequent evolution into a fully developed stall cell. Also, shown are the predictions of range extension with stall control and comparison to measurements. Finally, a status of the unsteady fullannulus simulations of the compression system components of the U.S. Army Helicopter Gas Turbine Engine is presented. All military and commercial gas turbine engine systems can benefit from the proposed work. In the future, this work could lead to new gas turbine engine designs with resistance to compressor stall and thus improved combat capability. This work will provide the first-ever 3-dimensional viscous time-accurate simulation of the stall inception process in a multistage compressor, providing insight into the causal link between compressor blade design parameters and the stable operating limit, which will be used to guide new design practices leading to compressor designs with increased tolerance to stall. This paper presents final-year progress on this challenge project.

Injection of a mixture of water/methanol is sometimes used to boost the power of a gas turbine engine, normally for short durations. The size of the droplets and their distribution at the point of injection in the inlet of the engine have a predominant effect on the engine performance. There is a tendency for larger droplets to migrate toward the casing of the compressor and impinge on it causing the casing to shrink, thereby reducing the blade clearances. Smaller droplets, injected near the hub surface, tend to follow the streamlines more closely and evaporate faster. The faster and complete evaporation of the droplets brings about an optimum decrease in air temperature, thereby improving the compressor performance. To aid in the design of an optimum water/methanol injection system, a threedimensional aerothermal analysis code was developed to study the droplet trajectory and evaporation characteristics in an axial flow compressor. The analysis was carried out on two multistage axial flow compressors. Good comparison was obtained between analysis and limited available test data. Nomenclature Pfs droplet mixture vapor pressure Mp molecular weight of evaporated droplet mixture P air pressure MA air molecular weight Ts droplet temperature Too ambient temperature Levap latent heat of vaporization TCr liquid critical temperature TEN liquid boiling temperature t time x axial coordinate r radial coordinate 0 angular coordinate * Associate Fellow of AIAA **MemberofAIAA Copyright © 1998 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. Vx axial velocity Vr radial velocity Ve angular velocity ^i absolute viscosity Prg Prandtl number of gas stream Cp specific heat p density d droplet diameter kg thermal conductivity of the gas Introduction Water-methanol injection into a gas turbine compressor has long been recognized as a method to increase the peak power that can be extracted from an engine. This technology can be adapted to boost the power of a turbo-shaft engine for a tiltrotor aircraft application. With this technology, for a given aircraft payload and operating range, the engine size can be significantly reduced if water-methanol injection is used to increase the maximum shaft horsepower. These higher power requirements are typically for short durations during the vertical takeoff and hovering modes of operation of the aircraft. The requirement of reducing the size of the engine is to be balanced with the amount of water-methanol to be carried on board as well as the additional system complexity that has to be introduced. A major problem associated with water injection into an axial flow compressor is that the injected droplets tend to impinge on the compressor casing, reducing the casing temperature and causing the casing to shrink. The shrinkage causes the compressor blades to gouge the casing, leaving a large gap that results in a permanent loss in compressor efficiency. To prevent this, compressors may be designed with large tip clearances. However, this would lower the compressor efficiency during noninjection operation regimes such as cruise of the aircraft. Many requirements must be met for a watermethanol injection system to be operational. First, the system must not allow injected water to impinge on the casing surface. Second, the water injection system should induce a high level of water-methanol evapora1 American Institute of Aeronautics and Astronautics tion to maximize the reduction in compressor air temperature and, therefore, maximize the increase in output power. In addition, the system should be simple, cost effective, low weight, and integrate well with the engine control system. A conceptual design of a watermethanol injection system for a civil tiltrotor application is shown in Figure 1. As a first step in the design of an efficient watermethanol injection system,we decided to develop a 3-D aerothermal analysis code to simulate water injection into a compressor. This code aids in the design of an optimized injection system for maximum evaporation of the droplets with minimal impingement on the casing. This report presents the experimental data on an engine test carried out with water-methanol injection, development of a three-dimensional aerothermal analysis code, and the results of the analysis and limited comparison with test data.

Gas turbine power enhancement technologies, such as inlet fogging, interstage water injection, saturation cooling, inlet chillers, and combustor injection, are being employed by end users without evaluating the potentially negative effects these devices may have on the operational integrity of the gas turbine. Particularly, the effect of these add-on devices, off-design operating conditions, nonstandard fuels, and compressor degradation∕fouling on the gas turbine’s axial compressor surge margin and aerodynamic stability is often overlooked. Nonetheless, compressor aerodynamic instabilities caused by these factors can be directly linked to blade high-cycle fatigue and subsequent catastrophic gas turbine failure; i.e., a careful analysis should always proceed the application of power enhancement devices, especially if the gas turbine is operated at extreme conditions, uses older internal parts that are degraded and weakened, or uses nonstandard fuels. This paper discusses a simplified method to evaluate the principal factors that affect the aerodynamic stability of a single-shaft gas turbine’s axial compressor. As an example, the method is applied to a frame-type gas turbine and results are presented. These results show that inlet cooling alone will not cause gas turbine aerodynamic instabilities, but that it can be a contributing factor if for other reasons the machine’s surge margin is already slim. The approach described herein can be employed to identify high-risk applications and bound the gas turbine operating regions to limit the risk of blade life reducing aerodynamic instability and potential catastrophic failure.

The study of the adhesion of micro sized particles to gas turbine internal surfaces, commonly known as gas turbine fouling, has gained increasing attention in the last years due to its dramatic effect on machine performance and reliability. On-field fouling analysis is mostly related to visual inspections during overhaul and/or programmed stops, which are performed, in particular, when gas turbine performance degradation falls under predetermined thresholds. However, these analyses, even if performed in the most complete as possible way, are rarely (or never) related to the conditions under which the gas turbine contamination takes place since the affecting parameters are difficult or even impossible to be adequately monitored. In the present work, a small scale multistage axial compressor is used to experimentally simulate the fouling phenomenon. The test rig allows the accurate control of the most relevant operating parameters which influence the fouling phenomenon. The compressor performance loss due to particle contamination has been quantitatively assessed. Soot particles appear stickier, especially in the presence of high humidity, and represent the most harmful operating conditions for the compressor unit. The deposits on the stator vanes and the rotor blades have been detected and post-processed, highlighting the most affected regions of each compressor stage employing an image analysis package tool.

In multi-stage compressors, the Variable Stator Vane (VSV) schedule significantly influences performance and stability. This study examines the part-span rotating stall in a multi-stage axial compressor which can lead to Non-Synchronised Vibration (NSV). Unsteady aerodynamic simulations were performed using URANS CFD for an axial compressor that experienced high vibration levels during the operation. For the original VSV schedule, steady-state simulations revealed the formation of short-length-scale disturbances near the tip of the first stage rotor. Full-annulus unsteady simulations demonstrated that these disturbances coalesce into rotating stall cells propagating around the annulus. The frequency of the unsteadiness caused by rotating structures was close to blade’s first mode natural frequency which could potentially cause NSV. Based on the findings, a new VSV schedule was developed, effectively eliminating the flow separation near the tip of the first-stage rotor and consequently rid the compressor of part-span rotating stall and NSV.

Multi-stage axial compressors play a crucial role in aerospace propulsion systems, as their exit flow field characteristics directly impact engine performance and stability. This study conducted numerical simulations on the first 3.5 stages of the NASA 74A transonic multi-stage axial compressor to analyze the exit flow field characteristics under different typical operating conditions. The research primarily investigated airflow deflection angle, radial velocity distribution, and their variation patterns. Additionally, the effects of inlet airflow angle and pressure variations on the exit flow field under non-uniform inlet conditions were examined in detail. The results indicate that at 68% rotational speed, the exit flow field of the NASA 74A compressor deteriorates significantly, with noticeable changes in distribution patterns. For the other four operating conditions, as the rotational speed decreases, both velocity and airflow angle exhibit a positive correlation with rotational speed. Compared to the design condition, peak velocity decreases by 2%, 3.7%, and 7%, while airflow deflection angle changes remain within 3°. Under non-uniform inlet conditions, when the inlet airflow angle decreases from 90° to 70°, variations in peak and average exit velocities remain within 2%, and the changes in peak and average airflow deflection angles are within 1%. However, when the inlet airflow angle decreases from 90° to 70°, the curve of the airflow deflection angle exhibits a leftward shift, with a deviation of 2.6%. Meanwhile, changes in inlet pressure under non-uniform conditions have a relatively minor impact on the overall flow field but significantly affect local distributions. When the inlet pressure increases from 1 atm to 1.05 atm, peak velocity increases by 0.98%, and average velocity rises by 3%. The maximum velocity difference reaches 6%, while the average airflow deflection angle differs by 0.7%, with a maximum deviation of 1.9°. Overall, the compressor exit flow field undergoes significant variations under different operating conditions, with increased flow instability at lower rotational speeds leading to flow separation, low-energy fluid accumulation, and non-uniform pressure distribution. In contrast, non-uniform inlet conditions have a relatively minor effect on the overall flow field but induce noticeable local changes, providing theoretical insights for compressor design optimization and performance evaluation.

The subject of this article is the impact of air delivery method bypassed from the compressor to the power turbine of a turboshaft gas turbine engine (GTE). The relevance of this GTE scheme is due to efforts to improve the efficiency of converting heat into work and to address the issue of ensuring stable compressor operation in throttling modes with a high compression ratio. A high compression ratio is necessary because increasing the cycle parameters, such as the heating degree and the pressure ratio of the working fluid, is essential for improving the overall efficiency of the GTE. One effective method to enhance the compressor stability is to bypass air from the compressor to the atmosphere, but this approach worsens the specific parameters of the GTE. In advanced GTE designs, various methods are considered to use bypassed air to avoid the loss of useful work. One such scheme is a turboshaft GTE with an air bypass to the power turbine. This design offers higher efficiency than bypassing air into the atmosphere while maintaining compressor stability. The GTE has an additional channel for delivering air from the compressor to the power turbine and a mixing chamber for the flows before the power turbine. The mixing chamber can receive either frontal or normal air delivery from the compressor. In the case of normal air delivery, the length of the mixing chamber is reduced, but this increases the total pressure losses and the resulting flow velocity before the power turbine. The aim of this work was to develop a mathematical model of a mixing chamber with normal flow delivery and to compare its characteristics with those of a mixing chamber with frontal flow delivery using an established mathematical model. To achieve this goal, the following tasks were accomplished: assuming no friction on the channel walls, an exact solution to the system of equations describing the mixing process was obtained. The mathematical relationships are presented in dimensionless form. The characteristics of mixing chambers with frontal and normal air delivery were obtained as dependencies of the resulting flow velocity and relative total pressure at the chamber outlet on the resulting flow velocity of the main stream and mass flow ratio of the streams, given the specified full temperature ratio of the streams at the mixing chamber inlet. A comparison was made between the flow parameters at the outlet of the mixing chambers with frontal and normal air delivery, under identical inlet conditions. The conclusions show that with normal flow delivery, the total pressure losses and the resulting velocity at the mixing chamber outlet are higher than with frontal flow delivery. However, the slight degradation in the flow parameters with normal flow delivery keeps such a mixing chamber competitive due to its shorter length.

The axial compressor of a gas turbine engine is the main part that determines the characteristics of the engine (aerodynamic characteristics, geometry, weight, efficiency, reliability, etc.). The technological features of blades in axial compressors have a fundamental influence on the aerodynamicdynamic characteristics of engines and their reliability during operation. The strength characteristics of the compressor blade and rotor also have a basic effect on the reliability performance; then, the breakage of the compressor blade can lead to local destruction of the engine and cause disaster for the plant. Considering the operating conditions of compressor blades and their role in gas turbine engines, it is found that compressor blades are a decisive factor in the reliability and quality of engines. When the compressor blade is damaged, it is understood as a change in geometrical characteristics, weight, physical and chemical properties of the substrate and coating, etc., due to the influence during engine operation. For the outside of the blade, the leading edge of the blade is often eroded (changing the radius of the leading edge, reducing the length of the pitch); damage by foreign bodies (significant deformation of the profile); change the shape of the tips of the blades; excessive roughness of the surface of the blades. It is mainly caused by foreign objects such as dust particles in the air, contamination of the blades, and significantly increased surface roughness of the blades. Therefore, it is necessary to pay attention to the exploitation of the engine and to solve the problem of causes and rules of erosion, pitting, and roughness. The article presents the results of studying the movement rules of dust particles in the clearance of the flow path of the axial compressor and the general characteristics of erosion, the change in shape of the compressor blade and other parameters and the characteristics of the gas turbine engine. This article presents a model to calculate the basic parameters of air flow and dust particles moving in the flow path of the compressor. Simulation can be performed using the ANSYS Fluent application, thereby comparing the assessment of the actual results with those of surveys. The results presented in the envelope are the basis for studying the negative effects of operating conditions on the engine, as well as proposing effective measures to protect the compressor blade from erosion and the methods of diagnostics of gas turbine engine.

This study presents experimental investigations into blade suction surface boundary layer flow in a multistage axial flow compressor. The experiments were focused on the third stage of the 4-stage Low Speed Research Compressor (LSRC) at Cranfield University. Measurements within the boundary layer were obtained using a hot wire probe. This was traversed normal to the blade surface at small increments, capturing the unsteady velocity profile within the boundary layer. Detailed boundary layer flow measurements covering most of the stator suction surface were taken and are described using time mean and ensemble averaged velocity profiles. Turbulence intensity in the boundary layer flow on the blade suction surface is also discussed. A strong wake-induced strip zone due to passing wake disturbances are generated at midspan near the blade leading edge at rotor blade passing frequency. Corner separation was observed at the tip region near the trailing edge. Normalized velocity profiles in this region show no variation in time.

Multistage compressor is the most important constituent of gas turbines used in land, naval and aeronautical applications. Overall performance of such machinery depends mainly on the axial compressor performance. Due to the relative motion between rotor and stator blades, the flow field in this machinery is highly unsteady. Furthermore several technological effects like tip clearances, complexity of the blade shapes, variation of axial distance between stator and rotor, seal leakages and cooling holes among others complicate the machine. Therefore the study of a complicated, strongly three-dimensional flow field inside a compressor is considered to be one of the most difficult tasks to be performed by a CFD expert. The present work is the extensive numerical study of the effect of: (1) tip clearance of rotor blades and (2) the axial gap between rotor and stator on the overall performance of a multistage axial compressor. A commercial software package is used for this study. Reynolds-averaged Navier–Stokes equations are solved using Spalart–Allmaras model. A number of steady-state viscous flow simulations were run for both the tip clearance effect and different axial gaps between stator and rotor. All simulations were performed for the first stage, i.e. Stator–Rotor–Stator. Simulations were carried out with coarse, medium and fine meshes to find an optimum, mesh-independent solution. It has been found that larger tip clearance has a detrimental effect on the stage pressure ratio and efficiency of a multistage axial compressor. Similarly there exists a certain distance ratio between the stator1–rotor and rotor–stator2, where stage performance is optimum. Overall performance characteristics obtained through simulation for both the tip clearance and axial gap variation were also compared with the experimental studies and found to be in good agreement.

Tip clearance has great influence on the performance of multistage axial compressors including efficiency, pressure rise, mass flow, as well as matching. This paper reports a study into the influence of tip clearance on the performance and matching of a 5-stage axial compressor by a numerical method. Different tip clearances from 0% to 5.0% span which represents the typical range of tip clearance in modern multistage axial compressors were simulated and analyzed. The results show that as tip clearance increases from 0% to 5.0% span, the choked mass flow decreases by about 21.8%, the peak pressure ratio decreases by about 43.1% and the peak efficiency decreases about 14.3 percents. As tip clearance increases, the efficiency of the whole compressor decreases in a parabolic manner not linearly as previous suggested, which is partially attributable to the cantilevered stators considered in this paper and primarily due to the mismatching of different stages. It is of great importance to control the tip clearance. When tip clearance increases, the front stage tends to work near surge condition and the rear stage tends to work near choke condition, which leads to lower efficiency than in the middle stages. A weight was defined to evaluate each stage’s contribution to the whole compressor’s efficiency deficit caused by the increase of tip clearance. Front and rear stages contribute more to the efficiency deficit than the middle stages, which indicates that more attention should be paid on front and rear stages to improve the performance of multistage axial compressors. In order to evaluate the matching of multistage axial compressors with a quantified method, a new parameter named “Peak Efficiency Deviation (PED)” was defined based on the difference between each stage’s operating efficiency and its peak efficiency. The mass flow of multistage axial compressors should be well considered to make the PED parameter to be close to zero as possible. In the most commonly used range of tip clearance from 0.5% to 3.0% span, the PED varies little within 0.4 percent, which is only about 8.4% of the peak efficiency deficit at 1.5% span tip clearance. So, the PED could be small within a wide range of tip clearances if the matching of the compressor is perfect at design tip clearance.

In this paper, an unsteady tip leakage flow instability is identified and investigated for an axial compressor at near surge conditions. We describe the results of experimental verification of a new compressor developed by improving the blade geometry based on the criterion for the occurrence of this unsteady phenomenon. In a high-speed multistage axial flow compressor having a subsonic high stagger rotor blade, a surge test was carried out by changing the tip clearance. Under a condition of large tip clearance, a drastic decrease in static pressure rise coefficient near the surge point was observed. At this operating condition, large, unsteady pressure fluctuation at the blade tip was confirmed, and the occurrence of tip leakage vortex breakdown that was clarified by unsteady multipoint pressure measurement and detailed unsteady numerical simulations. Due to the blockage effect caused by vortex breakdown of the tip leakage, double leakage and axially reversed flow near the trailing edge was observed. It is found that the vortex breakdown region of the tip leakage vortex propagated in the circumferential direction and caused the rotating instability. In order to investigate the relationship between this unsteady flow phenomenon, tip clearance size and flow pattern, unsteady calculation was conducted by changing the blade tip stagger and tip clearance size. A new concept of tip clearance of staggered pitch reference is proposed, which makes it possible to include the effect of blade loading on the clearance and clarifies that there exists a threshold at which vortex breakdown occurs/does not occur. On the basis of the above results, a high-speed multi-stage improved compressor was designed and manufactured to prevent tip leakage vortex breakdown. A clearance change test using active clearance control technology was conducted and increase in the static pressure rise coefficient near the surge point was confirmed for each clearance. The design concept of the improved blade which suppressed the unsteady tip leakage flow instability was tested and verified, and the effectiveness of the design guideline in actual gas turbines for power generation was confirmed.

In multistage axial compressors used for gas turbine engines, it is import ant to clarify unsteady phenomena such as surge and rotating stall. The purpose of this study is to develope a practical analysis method for both the surge and rotating stall in axial compressors. In this paper, firstly, some characteristics of the unsteady phenomena in axial compressors are summarized, derived from the measurments in test compressors. Secondary, a numerical method for the unsteady phenomena assuming incompressible flow is described. A feature of the method is solving the equations by Galerkin's method with the circumferential flow distortion expressed in the form of a high-order Fourier series. Finally, the usefulness of the method is examined by some results from a parameter-survey computation and by a comparison with the measured results in a 3-stage test compressor.