scholarly journals Cyclic Thermal Shock Damage Behavior in CVI SiC/SiC High-Pressure Turbine Twin Guide Vanes

Materials ◽  
2021 ◽  
Vol 14 (20) ◽  
pp. 6104
Author(s):  
Xiaochong Liu ◽  
Xiaojun Guo ◽  
Youliang Xu ◽  
Longbiao Li ◽  
Wang Zhu ◽  
...  
Keyword(s):  
High Pressure ◽  
Thermal Shock ◽  
Failure Modes ◽  
Guide Vane ◽  
Chemical Vapor ◽  
Leading Edge ◽  
Chemical Vapor Infiltration ◽  
Thermal Shock Test ◽  
High Pressure Turbine ◽  
Guide Vanes

In this paper, the SiC/SiC high-pressure turbine twin guide vanes were fabricated using the chemical vapor infiltration (CVI) method. Cyclic thermal shock tests at different target temperatures (i.e., 1400, 1450, and 1480 °C) in a gas environment were conducted to investigate the damage mechanisms and failure modes. During the thermal shock test, large spalling areas appeared on the leading edge and back region. After 400 thermal shock cycles, the spalling area of the coating at the basin and back region of the guide vane was more than 30%, and the whole guide vane turned gray, due to the formation of SiO2. When the thermal shock temperature increased from 1400 to 1450 and 1480 °C, the spalling area of the basin and the back region of the guide vane did not increase significantly, but the delamination occurred at the tenon, upper surface of the guide vane near the trailing edge of the guide vane. Through the X-ray Computed Tomography (XCT) analysis for the guide vanes before and after thermal shock, there was no obvious damage inside of guide vanes. The oxidation of SiC coating and the formation of SiO2 protects the internal fibers from oxidation and damage. Further investigation on the effect of thermal shock on the mechanical properties of SiC/SiC composites should be conducted in the future.

Comprehensive Geometric Description of Manufacturing Scatter of High-Pressure Turbine Nozzle Guide Vanes for Probabilistic CFD Analysis

2019 ◽  
Vol 141 (8) ◽  
Author(s):  
Paul Voigt ◽  
Lars Högner ◽  
Barbara Fiedler ◽  
Matthias Voigt ◽  
Ronald Mailach ◽  
...  
Keyword(s):  
High Pressure ◽  
Film Cooling ◽  
Guide Vane ◽  
Input Parameter ◽  
Three Dimensional ◽  
Leading Edge ◽  
Scanning System ◽  
High Pressure Turbine ◽  
Pressure Loss Coefficient ◽  
Nozzle Guide

The increasing demands on jet engines require progressive thermodynamic process parameters, which typically lead to higher aerothermal loadings and accordingly to designs with high complexity. State-of-the-art high-pressure turbine (HPT) nozzle guide vane (NGV) design involves vane profiles with three-dimensional features including a high amount of film cooling and profiled endwalls (PEWs). Typically, the specific mass flow, also called capacity, which governs the engine's operation, is set by the HPT NGV. Hence, geometric variations due to manufacturing scatter of the HPT NGV's passage can affect relevant aerodynamic quantities and the entire engine behavior. Within the traditional deterministic design approach, the influences of those geometric variations are covered by conservative assumptions and engineering experience. This paper addresses the consideration of variability due to the manufacturing of HPT NGVs through probabilistic CFD investigations. To establish a statistical database, 80 HPT NGVs are digitized with a high precision optical 3D scanning system to record the outer geometry. The vane profiles are parametrized by a section-based approach. For this purpose, traditional profile theory is combined with a novel method that enables the description of NGV profile variability taking the particular leading edge (LE) shape into account. Furthermore, the geometric variability of PEWs is incorporated by means of principle component analysis (PCA). On this basis, a probabilistic system assessment including a sensitivity analysis in terms of capacity and total pressure loss coefficient is realized. Sampling-based methods are applied to conduct a variety of 3D CFD simulations for a typical population of profile and endwall geometries. This probabilistic investigation using realistic input parameter distributions and correlations contributes to a robust NGV design in terms of relevant aerodynamic quantities.

R&D OF SICF/SIC COMPOSITE FOR FUSION REACTOR MATERIAL

2011 ◽  
Vol 25 (31) ◽  
pp. 4212-4215
Author(s):  
HAN-KI YOON ◽  
HO-JUN CHO ◽  
AKIRA KOHYAMA ◽  
TETSUJI NODA
Keyword(s):  
Thermal Shock ◽  
Fusion Reactor ◽  
Chemical Vapor ◽  
Chemical Vapor Infiltration ◽  
Thermal Shock Test ◽  
Reactor Material ◽  
System Pressure ◽  
Sic Composites ◽  
Thermal Shock Behavior ◽  
Purge Gas

During the operation of a fusion reactor, a certain type of transients could induce rapid cooldown of the reactor with relatively high or increasing system pressure. This induces a high tensile stress at the inner surface of the blanket, which is called the thermal shock. Test Blanket Module (TBM) has been designed to account for the performance of DEMO-relevant blanket. But, the TBM uses helium as coolant and tritium as purge gas, and FM steel, SiC f / SiC composites as a structural material. The SiC f / SiC composites with a SiC sub-layer/ PyC interlayers were therefore fabricated in order to improve the oxidation resistance. In the refractory field SiC f / SiC composites need to investigate residual stress and thermal shock behavior when high chemical stability and corrosion resistance are necessary. In this study, SiC f / SiC composite were fabricated by NITE process and the chemical vapor infiltration (CVI) method. SiC f / SiC composites have been examined and the detail analyses of the microstructure and the mechanical strength were conducted. Various pyrolytic carbons ( PyC ) with SiC interlayer coatings were applied to the composites for improvement of the mechanical properties at the interface.

Comprehensive Geometric Description of Manufacturing Scatter of High Pressure Turbine Nozzle Guide Vanes for Probabilistic CFD Analysis

2018 ◽  
Author(s):  
Paul Voigt ◽  
Lars Högner ◽  
Barbara Fiedler ◽  
Matthias Voigt ◽  
Ronald Mailach ◽  
...  
Keyword(s):  
High Pressure ◽  
Film Cooling ◽  
Guide Vane ◽  
Input Parameter ◽  
Three Dimensional ◽  
Leading Edge ◽  
Scanning System ◽  
High Pressure Turbine ◽  
Pressure Loss Coefficient ◽  
Nozzle Guide

The increasing demands on jet engines require progressive thermodynamic process parameters, which typically lead to higher aerothermal loadings and accordingly to designs with high complexity. State of the art high pressure turbine (HPT) nozzle guide vane (NGV) design involves vane profiles with three-dimensional features including a high amount of film cooling and profiled endwalls (PEW). Typically, the specific massflow, also called capacity, which governs the engine’s operation, is set by the HPT NGV. Hence, geometric variations due to manufacturing scatter of the HPT NGV’s passage can affect relevant aerodynamic quantities and the entire engine behavior. Within the traditional deterministic design approach, the influences of those geometric variations are covered by conservative assumptions and engineering experience. This paper addresses the consideration of variability due to manufacturing of HPT NGVs through probabilistic CFD investigations. In order to establish a statistical database, 80 HPT NGVs are digitized with a high precision optical 3D scanning system to record the outer geometry. The vane profiles are parametrized by a section based approach. For this purpose, traditional profile theory is combined with a novel method that enables the description of NGV profile variability taking the particular leading edge (LE) shape into account. Furthermore, the geometric variability of PEWs is incorporated by means of principle component analysis (PCA). On this basis, a probabilistic system assessment including a sensitivity analysis in terms of capacity and total pressure loss coefficient is realized. Sampling-based methods are applied in order to conduct a variety of 3D CFD simulations for a typical population of profile and endwall geometries. This probabilistic investigation using realistic input parameter distributions and correlations contributes to a robust NGV design in terms of relevant aerodynamic quantities.

scholarly journals Evaluation of High Pressure Turbine Blade Coatings on an LM2500 Rainbow Rotor

1995 ◽  
Author(s):  
B. Nagaraj ◽  
G. Katz ◽  
A. F. Maricocchi ◽  
M. Rosenzweig
Keyword(s):  
High Pressure ◽  
Turbine Blades ◽  
Leading Edge ◽  
Service Evaluation ◽  
High Pressure Turbine ◽  
Barrier Coating ◽  
Marine Propulsion ◽  
Corrosion Resistant Coatings ◽  
Stage 1 ◽  
Plasma Sprayed

Two LM2500 rainbow rotors with repaired stage 1 and stage 2 high pressure turbine blades are being assembled for marine propulsion service evaluation by the US Navy. The blades have seen between 5,000 and 15,000 hours of service in the Navy’s Fleets. A number of corrosion resistant coatings including plasma sprayed CoCrAlHf (bill of material), composite plated CoCrAlHf, platinum aluminide, aluminum silicide, and physical vapor deposited yttria stabilized zirconia thermal barrier coating (PVD TBC) will be evaluated in the rainbow rotor. This paper will discuss the advantages and microstructures of the various coatings. Composite plated CoCrAlHf, and PVD TBCs were recently service evaluated in an industrial LM2500 rainbow rotor for 10,500 hours. Both these coatings performed well, although the PVD TBC had local spallation at the leading edge. This paper will review the details of performance of these two coatings in the industrial LM2500 application.

Ice Crystal Icing Investigation on a Honeywell Uncertified Research Engine in an Altitude Simulation Icing Facility

2020 ◽  
Author(s):  
Ashlie B. Flegel
Keyword(s):  
Particle Size ◽  
Guide Vane ◽  
Leading Edge ◽  
Inlet Guide Vane ◽  
Inlet Temperature ◽  
Ice Crystal ◽  
Guide Vanes ◽  
Inlet Guide Vanes ◽  
The Core ◽  
Break Up

Abstract A Honeywell Uncertified Research Engine was exposed to various ice crystal conditions in the NASA Glenn Propulsion Systems Laboratory. Simulations using NASA’s 1D Icing Risk Analysis tool were used to determine potential inlet conditions that could lead to ice crystal accretion along the inlet of the core flowpath and into the high pressure compressor. These conditions were simulated in the facility to develop baseline conditions. Parameters were then varied to move or change accretion characteristics. Data were acquired at altitudes varying from 5 kft to 45 kft, at nominal ice particle Median Volumetric Diameters from 20 μm to 100 μm, and total water contents of 1 g/m3 to 12 g/m3. Engine and flight parameters such as fan speed, Mach number, and inlet temperature were also varied. The engine was instrumented with total temperature and pressure probes. Static pressure taps were installed at the leading edge of the fan stator, front frame hub, the shroud of the inlet guide vane, and first two rotors. Metal temperatures were acquired for the inlet guide vane and vane stators 1–2. In-situ measurements of the particle size distribution were acquired three meters upstream of the engine forward fan flange and one meter downstream of the fan in the bypass in order to study particle break-up behavior. Cameras were installed in the engine to capture ice accretions at the leading edge of the fan stator, splitter lip, and inlet guide vane. Additional measurements acquired but not discussed in this paper include: high speed pressure transducers installed at the trailing edge of the first stage rotor and light extinction probes used to acquire particle concentrations at the fan exit stator plane and at the inlet to the core and bypass. The goal of this study was to understand the key parameters of accretion, acquire particle break-up data aft of the fan, and generate a unique icing dataset for model and tool development. The work described in this paper focuses on the effect of particle break-up. It was found that there was significant particle break-up downstream of the fan in the bypass, especially with larger initial particle sizes. The metal temperatures on the inlet guide vanes and stators show a temperature increase with increasing particle size. Accretion behavior observed was very similar at the fan stator and splitter lip across all test cases. However at the inlet guide vanes, the accretion decreased with increasing particle size.

Computation of the Temperature Distribution in Cooled Radial Inflow Turbine Guide Vanes

1977 ◽  
Author(s):  
W. Tabakoff ◽  
W. Hosny ◽  
A. Hamed
Keyword(s):  
Temperature Distribution ◽  
Film Cooling ◽  
Guide Vane ◽  
Numerical Technique ◽  
Leading Edge ◽  
Guide Vanes ◽  
Heat Transfer Augmentation ◽  
Critical Areas ◽  
Internally Cooled ◽  
Convection Cooling

A two-dimensional finite-difference numerical technique is presented to determine the temperature distribution of an internally-cooled blade of radial turbine guide vanes. A simple convection cooling is assumed inside the guide vane. Such an arrangement results in relatively small cooling effectiveness at the leading edge and at the trailing edge. Heat transfer augmentation in these critical areas may be achieved by using impingement jets and film cooling. A computer program is written in Fortran IV for IBM 370/165 computer.

Using CFD to Reduce Resonant Stresses on a Single-Stage, High-Pressure Turbine Blade

2002 ◽  
Author(s):  
J. P. Clark ◽  
A. S. Aggarwala ◽  
M. A. Velonis ◽  
R. E. Gacek ◽  
S. S. Magge ◽  
...  
Keyword(s):  
High Pressure ◽  
Turbine Blade ◽  
Guide Vane ◽  
Turbine Blades ◽  
Suction Side ◽  
Lessons Learned ◽  
Single Stage ◽  
Navier Stokes ◽  
High Pressure Turbine ◽  
Time Resolved

The ability to predict levels of unsteady forcing on high-pressure turbine blades is critical to avoid high-cycle fatigue failures. In this study, 3D time-resolved computational fluid dynamics is used within the design cycle to predict accurately the levels of unsteady forcing on a single-stage high-pressure turbine blade. Further, nozzle-guide-vane geometry changes including asymmetric circumferential spacing and suction-side modification are considered and rigorously analyzed to reduce levels of unsteady blade forcing. The latter is ultimately implemented in a development engine, and it is shown successfully to reduce resonant stresses on the blade. This investigation builds upon data that was recently obtained in a full-scale, transonic turbine rig to validate a Reynolds-Averaged Navier-Stokes (RANS) flow solver for the prediction of both the magnitude and phase of unsteady forcing in a single-stage HPT and the lessons learned in that study.

Assessment of the Indirect Combustion Noise Generated in a Transonic High-Pressure Turbine Stage

2015 ◽  
Author(s):  
Dimitrios Papadogiannis ◽  
Florent Duchaine ◽  
Laurent Gicquel ◽  
Gaofeng Wang ◽  
Stéphane Moreau ◽  
...  
Keyword(s):  
High Pressure ◽  
Acoustic Waves ◽  
Gas Turbines ◽  
Guide Vane ◽  
Combustion Noise ◽  
Dynamic Mode Decomposition ◽  
Dynamic Mode ◽  
Noise Generation ◽  
Turbine Stage ◽  
High Pressure Turbine

Indirect combustion noise, generated by the acceleration and distortion of entropy waves through the turbine stages, has been shown to be the dominant noise source of gas turbines at low-frequencies and to impact the thermoacoustic behavior of the combustor. In the present work, indirect combustion noise generation is evaluated in the realistic, fully 3D transonic high-pressure turbine stage MT1 using Large-Eddy Simulations (LES). An analysis of the basic flow and the different turbine noise generation mechanisms is performed for two configurations: one with a steady inflow and a second with a pulsed inlet, where a plane entropy wave train at a given frequency is injected before propagating across the stage generating indirect noise. The noise is evaluated through the Dynamic Mode Decomposition of the flow field. It is compared with previous 2D simulations of a similar stator/rotor configuration, as well as with the compact theory of Cumpsty and Marble. Results show that the upstream propagating entropy noise is reduced due to the choked turbine nozzle guide vane. Downstream acoustic waves are found to be of similar strength to the 2D case, highlighting the potential impact of indirect combustion noise on the overall noise signature of the engine.

Numerical Prediction of Cooling Performance Sensitivity of 1st Stage Nozzle Guide Vane Under Aerothermal Conditions

2019 ◽  
Author(s):  
Prasert Prapamonthon ◽  
Bo Yin ◽  
Guowei Yang ◽  
Mohan Zhang
Keyword(s):  
Heat Transfer ◽  
Surface Roughness ◽  
High Pressure ◽  
Guide Vane ◽  
Pressure Ratio ◽  
Inlet Temperature ◽  
High Pressure Turbine ◽  
Cooling Performance ◽  
Nozzle Guide Vane ◽  
Nozzle Guide

Abstract To obtain high power and thermal efficiency, the 1st stage nozzle guide vanes of a high-pressure turbine need to operate under serious circumstances from burned gas coming out of combustors. This leads to vane suffering from effects of high thermal load, high pressure and turbulence, including flow-separated transition. Therefore, it is necessary to improve vane cooling performance under complex flow and heat transfer phenomena caused by the integration of these effects. In fact, these effects on a high-pressure turbine vane are controlled by several factors such as turbine inlet temperature, pressure ratio, turbulence intensity and length scale, vane curvature and surface roughness. Furthermore, if the vane is cooled by film cooling, hole configuration and blowing ratio are important factors too. These factors can change the aerothermal conditions of the vane operation. The present work aims to numerically predict sensitivity of cooling performances of the 1st stage nozzle guide vane under aerodynamic and thermal variations caused by three parameters i.e. pressure ratio, coolant inlet temperature and height of vane surface roughness using Computational Fluid Dynamics (CFD) with Conjugate Heat Transfer (CHT) approach. Numerical results show that the coolant inlet temperature and the vane surface roughness parameters have significant effects on the vane temperature, thereby affecting the vane cooling performances significantly and sensitively.

Sign in / Sign up

Export Citation Format

Share Document