Rotor Blade Pressure Research Articles

Study on erosion wear characteristics of aero-compressor blades considering distortion degree

A Tabakoff model of a certain type of turboshaft engine compressor blade was established, and the effects of no, mild, moderate, and severe distortions on blade erosion pattern were obtained. Based on particle velocimetry and erosive wear experiments, the Tabakoff model parameters (k12, V1, V2, V3, and θ0) were obtained. The effect of four distortion degrees on blade erosion pattern and erosion rate was studied. With an increase in the distortion degree, the erosion area of the rotor blade pressure surface is found to move toward the leading edge and blade tip. The erosion area of the stator blade suction surface moves toward the leading edge and blade root, whereas that of the pressure surface moves toward the blade root.

Analysis of the effects of simplifications on the state of loads in a centrifugal compressor

This study aims to quantify and assess qualitatively the impact of modeling simplifications used to represent inertial and aerodynamic loads on the stresses and structural deformations of a centrifugal compressor in operation. The research object is the compressor of the high-pressure line of the DGEN 380 bypass turbine engine. Based on the virtual dynamometer WESTT CS/BV, the gas-dynamic parameters at the entrance to the centrifugal compressor and after the stage are determined. These values were used as initial parameters for numerical flow analysis. As part of the numerical strength analyses, a series of several load configurations were carried out: spin only, spin and inlet pressure normally applied on the working surface of the rotor blade, spin and outlet pressure normally applied on the working surface of the rotor blade, and one-way fluid–structure interaction analysis taking into account the aerodynamic loads with and without spinning. Based on the simulations, the level of similarity of a given load configuration with the last analysis, adopted as the reference, was determined. It was observed that in terms of the stress state, the rotational analysis taking into account the pressure on both sides of the blade gives satisfactory results, but the strain values are overestimated. The results obtained and the method of evaluation of compressor results can be used in research in the area of aviation, automotive, and refrigeration industries.

A NOVEL DESIGN OF PERISTALTIC CARBON NANO PUMP AND AN ANALYSIS OF HELIUM FLOW

A novel nano-scale pump that can transport atoms or small molecules with a peristaltic motion is designed. It is proven by molecular-dynamics simulations that the introduced nano-pump design works properly. The designed nano-pump consists of one main carbon nanotube named the flow tube and two rotors where multi-walled carbon nanotubes are attached. The pumping of helium atoms by the designed peristaltic carbon nano-pump is investigated by molecular-dynamics simulations. For varying rotor speeds and blade counts, time-averaged velocity, temperature, and pressure results of pumped helium atoms are calculated, and relationships between them are modeled as polynomial surfaces. The results showed that rotor frequency increases the velocity of helium linearly and the temperature and pressure of helium non-linearly. Furthermore, the blade count of the proposed mechanism does not substantially affect the velocity as per the previous studies in the literature.

On the Interactions of a Rotor Blade Tip Flow With Axial Casing Grooves in an Axial Compressor Near the Best Efficiency Point

Experiments in a refractive index-matched axial turbomachine facility show that semicircular skewed axial casing grooves (ACGs) reduce the stall flowrate by 40% but cause a 2.4% decrease in the maximum efficiency. Aiming to elucidate mechanism that might cause the reduced efficiency, stereo-PIV measurements examine the impact of the ACGs on the flow structure and turbulence in the tip region near the best efficiency point (BEP), and compare them to those occurring without grooves and at low flowrates. Results show that the periodic inflow into the groove peaks when the rotor blade pressure side (PS) overlaps with the downstream end of the groove, but diminishes when this end faces the suction side (SS). Entrainment of the PS boundary layer and its vorticity generates a vortical loop at the entrance to the groove, and a “discontinuity” in the tip leakage vortex (TLV) trajectory. During exposure to the SS, the backward tip leakage flow separates at the entrance to the groove, generating a counter-rotating circumferential “corner vortex,” which the TLV entrains into the passage at high flowrates. Interactions among these structures enlarge the TLV and create a broad area with secondary flows and elevated turbulence near the groove's downstream corner. A growing shear layer with weaker turbulence also originates from the upstream corner. The groove also increases the flow angle upstream of the blade tip and varies it periodically. Accordingly, the circulation shed from the blade tip and strength of leakage flow increase near the blade leading edge (LE).

Effect of pyrometer type and wavelength selection on temperature measurement errors for turbine blades

Turbine blade temperature measurements are important for monitoring the working state of the blades. However, the reflected radiation from the high temperature surrounding environment leads to significant error in the optical pyrometer measurements. This study calculates and compares temperature error of single wavelength pyrometer, ratio pyrometer and multicolor pyrometer in measuring turbine rotor blade temperature. Emissivity of the blade was measured in the wavelength range of 1.2 μm to 2.5 μm. Temperature distribution of the rotor blade and the guide vane was simulated by CFD software. Additionally, temperature error was calculated based on - discretization of a three-dimensional blade model. The results show that from the leading edge to the trailing edge of the rotor blade pressure side, the variation trend of the three kinds of pyrometer temperature error is the same, first decreasing, then increasing before decreasing again. The maximum relative temperature error was found to be 7.5%, 4.9% and 2.6% when the wavelength of the pyrometer was 1.3 μm, 1.6 μm, 2.2 μm, respectively. The study also shows that wavelength selection has great influence on the ratio pyrometer. It was found that the maximum relative temperature error ranged from 2.6% to 15.7% based on the choice of the various wavelengths. On the other hand, the maximum relative temperature error of the multicolor pyrometer was 3.8%. The analysis presented in this work will be of importance in providing guidelines for choosing the optimum measurement wavelengths for optical pyrometers. Furthermore, the finding will also useful in the calculation and correction of the optical pyrometer error.

Experimental and numerical investigation on the pressure pulsation and instantaneous flow structure in a nuclear reactor coolant pump

Abstract The instantaneous vortical flow structure is one of the typical flow structures inside the nuclear reactor coolant pump (RCP), which would cause the unsteady pressure pulsations, vibrations of the unit and fatigue of components. Due to high safety requirement of the RCP during the actual operation of the nuclear power plants (NPPs), revealing instantaneous nature of vortical flow structure and its pressure pulsation becomes a crucial issue for studying the internal flow mechanism of the RCP. The purpose of this study is to shed comprehensive light on the pressure pulsation and instantaneous vortical flow structure in the RCP by using the experimental method and the numerical simulation method. Based on the result of comparison with the experiment and numerical simulation, it can be considered that the LES method can better identify instantaneous nature of vortical flow structures in the low frequency band. The instantaneous vortical flow structure is attached to the impeller RBPS (rotor blade pressure surface) and transient jet wake vortex flow structures between the impeller and diffuser under the rotor-stator interaction effect are fairly obvious by Q-criterion. On the other hand, in the axial-vorticity component distribution, the rotor-stator interaction effect between the impeller and diffuser cannot be clearly revealed. From the three-dimensional structures, the main vortex structures are in the front and back cavity of the right-hand side near the discharge nozzle and the right-hand side below the discharge nozzle in the spherical casing. Meanwhile, combined with pressure spectrum, it is convinced that their unsteady characteristic is the main reason that causes complex excitation frequencies in the low frequency band of the right-hand side near the discharge nozzle.

고속압력감응페인트를 이용한 로터 블레이드 표면 압력 측정

본 논문은 고속압력감응페인트기법(Fast-Responding PSP)을 소개하고, 이를 이용하여 정지비행상태에 있는 축소형 로터 블레이드의 표면(윗면) 압력을 측정해봄으로서 PSP를 이용한 로터 블레이드 표면 압력 측정의 정확성과 그에 따른 실험기법의 타당성을 검증하기 위하여 수행되었다. 실험을 위한 광원으로는 532 nm 파장을 가지는 Pulsed laser를 사용하였고, PSP 측정 기법으로는 Lifetime 기법을 적용하였다. 또한, 모델 표면에 도포된 압력감응페인트는 반응성이 높은 Porous PSP가 사용되었다. 로터 블레이드는 NACA0012 익형을 가지고 있으며 길이 340 mm, 코드 40 mm의 직사각 형상 1종과 끝단의 후퇴각이 다른 4종의 형상을 사용하였다. 로터 블레이드의 콜렉티브 피치각 변화에 따른 표면 압력 분포를 측정하였으며 측정된 결과를 통해 콜렉티브 피치각이 증가할수록 윗면의 압력이 낮아지는 것을 정성적으로 확인하였고, 정량적인 압력계수는 NASA의 실험 데이터와 비교하여 약 0.4~0.7 정도 높은 경향성을 보였다. The present study was conducted by using fast-responding PSP technique to measure the surface pressure on a small-scale rotor blade in hover. Also, the study was performed to verify the accuracy and investigate its possibility of PSP application for rotor blade pressure measurement. Pulsed laser which has 532 nm wavelength was used as a light source. Lifetime measurement technique was applied. Also, the coated paint on a rotor blade was porous PSP which has faster response time than conventional PSP. The blades had NACA0012 airfoils. The length of rotor blade was 340 mm and chord was 40 mm with rectangular shape 1 set, and 4 sets had several tip sweepback angles. The measured results qualitatively showed that the upper surface pressure decreases with increasing the collective pitch angle. Quantitative pressure coefficients of PSP results were higher approximately 0.4 to 0.7 than the pressure tap data of the NASA experiment.

Periodical Unsteady Flow Within a Rotor Blade Row of an Axial Compressor—Part I: Flow Field at Midspan

Abstract In this two-part paper, results of the periodical unsteady flow field within the third rotor blade row of the four-stage Dresden low-speed research compressor are presented. The main part of the experimental investigations was performed using laser Doppler anemometry. Results of the flow field at several spanwise positions between midspan and rotor blade tip will be discussed. In addition, time-resolving pressure sensors at midspan of the rotor blades provide information about the unsteady profile pressure distribution. In Part I of the paper, the flow field at midspan of the rotor blade row will be discussed. Different aspects of the blade row interaction process are considered for the design point and an operating point near the stability limit. The periodical unsteady blade-to-blade velocity field is dominated by the incoming stator wakes, while the potential effect of the stator blades is of minor influence. The inherent vortex structures and the negative jet effect, which is coupled to the wake appearance, are clearly resolved. Furthermore the time-resolved profile pressure distribution of the rotor blades is discussed. Although the negative jet effect within the rotor blade passage is very pronounced, the rotor blade pressure distribution is nearly independent of the convectively propagating chopped stator wakes.

Effect of Unsteady Stator Wake—Rotor Double-Leakage Tip Clearance Flow Interaction on Time-Average Compressor Performance

A study has been conducted, using unsteady three-dimensional Reynolds-averaged Navier-Stokes simulations to determine the impact on rotor performance of the interaction between upstream (steady defect and time-varying defect) stator wakes and rotor tip clearance flow. The key effects of the interaction between steady stator wakes and rotor tip clearance flow are: 1) a decrease in loss and blockage associated with tip clearance flow; 2) an increase in passage static pressure rise. Performance benefit is seen in the operability range from near design to high loading. The benefit is modest near design and increases with loading. Significant beneficial changes due to the stator-rotor interaction occur when the phenomenon of tip clearance flow double-leakage is present. Double-leakage occurs when the tip clearance flow passes through the tip gap of the adjacent blade. It is detrimental for compressor performance. The effect of strong stator-rotor interaction is to suppress double-leakage on a time-average basis. Double-leakage typically takes place at high loading but can be present at design condition as well, for modern highly loaded compressor. A benefit due to unsteady interaction is also observed in the operability range of the rotor. A new generic causal mechanism is proposed to explain the observed changes in performance. It identifies the interaction between the tip clearance flow and the pressure pulses, induced on the rotor blade pressure surface by the upstream wakes, as the cause for the observed effects. The direct effect of the interaction is a decrease in the time-average double-leakage flow through the tip clearance gap so that the stream-wise defect of the exiting tip flow is lower with respect to the main flow. A lower defect leads to a decrease in loss and blockage generation and hence an enhanced performance compared to that in the steady situation. The performance benefits increase monotonically with loading and scale linearly with upstream wake velocity defect. With oscillating defect stator wakes, rotor performance shows dependence on oscillation frequency. Changes in the tip region occur at a particular reduced frequency leading to (1) decrease in blockage, and (2) increase in passage loss. The changes in rotor performance at a particular reduced frequency are hypothesized to be associated with the inherent unsteadiness of the tip clearance vortex and its resonance behavior excited by the oscillating wakes.

Cycle Performance of a Combined Plant Using a Closed Circuit Steam Cooled Gas Turbine.

Closed circuit steam cooled gas turbines are drawing attention because of their efficiency compared with that of a conventional air cooled gas turbine. In the closed circuit steam cooled gas turbine, coolant steam is not discharged into the gas path, so dilution of the hot gas stream, rotor blade pumping loss and pressure loss due to mixing of coolant steam are drastically reduced. In this paper, combined cycle performance using closed circuit steam cooled gas turbine are analyzed. The evaluated parameters are bleed and return point of the coolant steam in the bottoming cycle and leakage, pressure loss, mass flow of the coolant steam. The result of this analysis is that the combined cycle efficiency reaches about 53.6% on the HHV bases.

Inlet Distortion Generated Periodic Aerodynamic Rotor Response

Fundamental inlet distortion-generated rotor blade row unsteady aerodynamics, including the effects of both the detailed aerodynamic forcing function for the first time and steady loading are experimentally investigated in an extensively instrumented axial-flow research compressor. A two-per-rev forcing function with three gust amplitude ratios is generated. On the rotor blade pressure surface, the unsteady pressure nondimensionalization compresses the magnitude data with mean flow incidence angle. This is not the case on the higher camber suction surface. These pressure surface unsteady data are primarily affected by the steady loading level, whereas the suction surface unsteady data are a function of the steady loading level and distribution as well as the gust amplitude ratio. In addition, a design inlet distortion blade surface unsteady pressure correlation is considered.

Measurement and Analyses of Heat Flux Data in a Turbine Stage: Part II—Discussion of Results and Comparison With Predictions

Parts I and II of this paper report the measurement and analysis of detailed heat flux distributions obtained on the nozzle guide vane (NGV), rotor, and shroud of the full-stage rotating turbine of the Garrett TFE 731-2 engine. Part I is devoted to a description of the experimental apparatus and the data analysis procedure. Part II is concerned with the experimental results and their comparison with predictions obtained using a flat-plate technique and with predictions obtained using STAN5. Measurements were performed for values of the ratio of wall temperature to total temperature equal to 0.21, 0.33, and 0.53. For the NGV airfoil and rotor blade, the influence of wall temperature ratio on the Stanton number is shown to be generally 10 percent or less. For the NGV airfoil, and the rotor blade pressure surface, the flat-plate turbulent-flow prediction provides a reasonably good comparison with the data. With the rotor operating, the state of the boundary layer generally appears to be turbulent for both the NGV airfoil and the rotor blade. Predictions obtained using STAN5 confirm that boundary layer transition occurs at a value of Reθ substantially less than 200.